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AD-762 198

EXPENDABLE MAIN ROTOR BLADE STUDY
John A. Longobardi, et al
United Aircraft Corporation

Prepared for:

Army Air Mobility Research and Development
Laboratory

April 1973

DISTRIBUTED BY:

National Technical Information Service
U. S. DEPARTMENT OF COMMERCE

5285 Port Royal Road, Springfield Va. 22151

5 USAAMRDL TECHNICAL REPORT 72-47

£>

EXPENDABLE MAIN ROTOR BLADE STUDY

John A. Longobardi
Everett Fournier

April 1973

EUSTIS DIRECTORATE

U. S. ARMY AIR MOBILITY RESEARCH AND DEVELOPMENT LABORATORY

Reproduced by

NATIONAL TECHNICAL
INFORMATION SERVICE

U 5 Deportment of Commerce
Springfield VA ??M]

FORT EUSTIS, VIRGINIA

CONTRACT DAAJ02-7T-C-0C46
SIKORSKY AIRCRAFT

DIVISION OF UNITED AIRCRAFT CORPORATION
STRATFORD, CONNECTICUT

Approved for public release;
distribution unlimited.

DISCLAIMERS

The findings in this report are not to be construed as an official Department of the Army
position unless so designated by other authorized documents.

When Government drawings, specifications, or other data are used for any purpose other
than in connection with a definitely related Government procurement operation, the
United States Government thereby incurs no responsibility nor any obligation whatsoever;
and the fact that the Government may have formulated, furnished, or in any way supplied
the said drawings, specifications, or other data ia not to be regarded by implication or
otherwise as in any manner licensing the holder or any other person or corporation, O’
conveying any rights or permission, to manufacture, use, or sell any patented invention
that may in any way be related thereto.

Trade names cited in this report do not constitute an official endorsement or approval of
the use of such commercial hardware or software.

DISPOSITION INSTRUCTIONS

Destroy this report when no longer needed. Do not return it to the originator.

IT.

DBTRIIUTIM/AMIUtlllTT CODES

Blit. AVAIL, uid or SPECIAL

Unclassified

Sri uritx (. I*issific.ilit»n

DOCUMENT CONTROL DATA ■ R & D

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1 ON'OINA TINL Activity fCuffloM/P ,ll/l/luf)

Sikorsky Aircraft
D.vision of United Aircraft
Stratford. Connecticut _

HI-POHT Sf. CUr<l TV CLASSIFICATION

Unclassified

ih. CROUP

K F POR T TITLE

EXPENDABLE MAIN ROTOR BLADE STUDY

4 Ut.SCRlPTlvE NOTE5f7)7J«‘ ol report .1 m/ inc/ijwvr ifiifc.v)

Einal-Separt.

Au ThORiSI tFirst nam «■, initial. Inst ii.mu-

John A. Longobardi
Everett Fournier

£• M t FOR T • & ’ r

April 1973

70. TOTAL "" !■ AGES

Ih NO OF BUIS

_18

H«l. CON TRAC T OR GHAh T NO

DAAJ02-71-C-0046

b. I'ROJFC T NO

1F1622205A11901

•i.

ORIGINATOR'S Hd'OIvi f.UMI'l RlS)

USAAMRDL Technical Report 72-47

iT OTHER REPORT -1013) t Anv uthvt mimhvrs that may In' tissifitivtl
this report)

Sikorsky Engineering Report 50748

ID DISTRIBUTION 5TATFMENT

Approved for public release; distribution unlimited.

IT SUPPLEMENTARY NOTI5

• i SPONSORING MIL! I AN V ACT Y I T V

Eustis Directorate, U.S. Army Air
Mobility Research and Development
Laboratory, Fort Eustis, Virginia

n A MS TRACT

This report presents Sikorsky's study of expendable blade designs applicable to the
Army's UH-1H helicopter with its teetering rotor system. The program included
design, reliability, maintainability and cost analysis studies. Reliability and
maintainability parameters were developed /hich were subsequently inserted into
cost model equations to determine life cycle cost comparisons of the new blade
designs with the present UH-1H blade.

More than fifteen configurations were investigated and reduced to six viable blade
designs. They included aluminum, steel, and composite configurations. The study
covered two time frames: 1972 and 1980. The results showed that a low-cost
aluminum extrusion (Configuration I) with a fiberglass composite skin was the most
cost effective for 1972. The 1980 time frame showed that an all-composite blade
(Configuration V) was the most cost effective.

The report also includes field repair procedures for the leading candidate blades
developed. A simulated field repair was performed demonstrating the feasibility
of composite/honeycomb repair. Also included is a future plan for hardware evalua¬
tion outlining the major phases in a development program for the most cost effect¬
ive blade for the 1980 time frame.

DD F .rj473

Unclassified

Security Classification

UH-1H Helicopter
Expendable Blades
Repairable Blades
Cost Effective Blades
Field Repairable Blades
Semirigid Teetering Rotor System
High Modulus Composite Blades
Automation of Blade Production

DEPARTMENT OF THE ARMY

U. S. ARMY AIR MOBILITY RESEARCH A DEVELOPMENT LABORATORY
EUBTIS DIRECTORATE
FORT EUSTI3, VIRGINIA 23004

This is one of a number of parallel studies examining various rotor
blade design concepts emphasizing reliability and maintainability.

Other concepts that have been studied are repairable and sectlonalized
rotor blade designs. A parallel expendable rotor blade study has been
performed by Kaman Aerospace Corporation. These design studies are
aimed at achieving considerable improvement In rotor blade R&M charac¬
teristics, thereby reducing life-cycle cost. To achieve comparability,
all blade designs are required to match UH-1D/H characteristics, and
life-cycle cost is compared to that for the UH-1D/H.

This study concentrated on designing a low-cost rotor blade that is
more cost effective to scrap than to return for depot level repair.

For the 1972 time frame, a blade with an aluminum extruded spar and
1 jneycomb-filled fiberglass afterbody was the most cost effective
configuration considered. Because of the predicted trends ■ material
and labor costs together with the anticipated automated p. sses for
composites, an all-composite configuration with a fibergla spar and
preformed carbon and fiberglass afterbody was projected to be the most
cost effective for the 1980 time frame.

The cost results, although valid for comparative purposes, cannot be
corsid'iud on an absolute scale. The blade design selected and the
repair procedures arrived at in this study T-.aSt also be tested under
operational conditions, as must the structural integrity of the repaired
blade.

The coi._-jrion that field-expendable rotor blade designs, as presented
in this Phase I report, are cost effective is supported by the results
of the parallel design study, although a different design approach was
selected. A Phase II report with comparative radar cross-section mea¬
surements for simulated Configurations I, IV, and UH-1 rotor blades is
in preparation. The results of this study and other related efforts
are being considered in a recently initiated procurement for the design
and development of a field-repairable/expendable rotor blade concept.

The program was conducted under the technical management of Philip J.
Haselbauer, Technology Applications Division, with engineering support
from Joseph H. McGarvey, Military Operations Technology Division.

I*.

KWfreom rw» *~»ew>.

Task 1F162205A11901
Contract DAAJ02-71-C-0046
USAAMRDL Technical Report 72-47
April 1973

EXPENDABLE MAIN ROTOR BLADE STUDY
Final Report - Phase I

Sikorsky Engineering Report 50748

John A. Longobardi
Everett Fournier

Prepared by

Sikorsky Aircraft

Division of United Aircraft Corporation
Stratford, Connecticut

for

EUSTIS DIRECTORATE

U.S. ARMY AIR MOBILITY RESEARCH AND DEVELOPMENT LABORATORY

FORT EUSTIS, VIRGINIA

Approved for public release;
distribution unlimited.

SUMMARY

The report presents results of a design study of expendable main rotor
blades for the UH-1H helicopter. The objective of the study was to de¬
sign blades which could eventually be thrown away after extensive damage
rather than be sent back to depot for major overhaul. Unit cost, field re-
payability, resistance to corrosion and erosion, fatigue strength, and
damage tolerance were factors considered for maximum cost effective¬
ness or lowest life-cycle cost. The study was limited to the UH-1 blade,
which requires a structural skin for edgewise rigidity. For an articu¬
lated rotor blade, some of the conclusions regarding skin construction
and material could be different.

The study included development of reliability, maintainability,and cost-
effectiveness models. In addition, the United Aircraft Normal Modes
Computer Program was modified to include two-bladed teetering rotor
dynamics. The cost model was based upon the present UH-1H aircraft to
provide life-cycle cost comparisons with the new blades designed in this
study.

More than fifteen blade designs were generated. They included alumi¬
num, steel, and composite blade designs. The study covered two time
frames: 1972 and 1980. The results showed that a low-cost aluminum
extrusion with a fiberglass composite skin is the most cost effective
blade for the 1972 time frame. This blade, which has 30% fewer parts,
was estimated to be 20% cheaper and 75% more repairable than the Bell
blade. It is estimated that this blade could save $12 million for a base¬
line fleet of 1,000 aircraft.

For the 1980 time frame, the Sikorsky "twin beam” all-composite blade
has the potential of being twice as repairable as the Bell blade and could
save $26 million for a baseline fleet of 1,000 aircraft. To realize this
potential, the costs of carbon must be reduced to $25 per pound. The
technology to automate the manufacture of the blade in one or two pieces
must be developed and demonstrated, and the ability to repair the spar
and trailing edge with sufficient remaining strength must also be demon¬
strated. Because of the potential of the twin beam concept and the re¬
search and development needed to demonstrate its production suitability,
the plan for hardware development is centered on development of this
concept.

FOREWORD

This design study for expendable main rotor blades was performed under
Contract DAAJ02-71-C-0046 with the Eustis Directorate, U. S. Army
Air Mobility Research ■ 1 Development Laboratory, Ft. Eustis, Virginia,
Task 1F162205A11901, J was under the general technical direction of
Mr. Philip J. Haselbauer of the Structures Division of USAAMRDL. This
expendable main rotor blade study is one of two cuch studies conducted as
follow-on studies to earlier sectionalized and repairable rotor blade ad¬
vanced design studies. The objective of all these studies was to obtain
more cost-effective blade concepts for Army utilization.

Sikorsky's principal participants were Everett F. Fournier of the Relia¬
bility and Maintainability Section, Mario J. D'Onofrio and John R. Olson
from the System Analysis Section, and William C. Reinfelder of the
Rotor System Section. John A. Longobardi, also from the Rotor System
Section, was the Team Task Manager. The program was under the gen¬
eral supervision of William F. Paul, Rotor System Section Head.

Preceding page blank

TABLE OF CONTENTS

SUMMARY.

FOREWORD.

LIST OF ILLUSTRATIONS.

LIST OF TABLES.

LIST OF SYMBOLS.

INTRODUCTION.

DEVELOPMENT OF METHODOLOGY AND DESIGN
CONFIGURATIONS.

ANALYSIS OF DESIGN CONFIGURATIONS .

DESIGN SELECTION.

CONCLUSIONS ..

RECOMMENDATIONS.

LITERATURE CITED .

APPENDIX I - BLADE CHARACTERISTICS.

APPENDIX II - RELIABILITY/MAINTAINABILITY DATA

APPENDIX III - COST-EFFECTIVENESS MODEL.

APPENDIX IV - UH-1H BLADE DATA.

APPENDIX V - COST EFFECTIVE COMPARISON USING
MTBR OF 1063 HOURS .

APPENDIX VI - PLAN FOR FUTURE HARDWARE

EVALUATION .

DISTRIBUTION .

Page

iii

viii
xiii
xv i

144

149

151

152
154
158
215
234

236

254

273

vii

Preceding page blank..

AXt-

LIST OF ILLUSTRATIONS

Figure Page

1 Impact of Nonrecurring, Shop Hours, Material Cost

on Blade Costs. 7

2 Cost-Effectiveness Model. 9

3 Life-Cycle Blade Logistics. 11

4 Material Cost Comparisons. 19

5 Configuration I.21

6 "C" Spar Blade.24

7 "C" Spar Blade With Backwall Channel . 24

8 UH-1 Root End.28

9 Aluminum Laminates Root End.28

10 Stepped Extrusion Root End.28

11 Solid Aluminum Root End.29

12 Fiberglass Laminates Root End.29

13 Reduced Doubler Root End.29

14 Roll-Formed Schematic. 31

15 Roll-Formed Spar. 32

16 Configuration II. 33

17 Configuration III.37

18 Flatwise and Torsional Stiffness Change with Spar

Chord Change.39

19 Configuration IV.41

20 Fabrication of Beam Concept.43

viii

Figure Page

21 Fiberglass Root End Attachment. 45

22 Configuration V . 47

23 Pultrusion Schematic. 49

24 Configuration VI. 51

25 Natural Frequencies, Configuration UH-1H. 54

26 Natural Frequencies, Configuration I and VI. 55

27 Natural Frequencies, Configuration II. 56

28 Natural Frequencies, Configuration III. 57

29 Natural Frequencies, Configuration IV and V. 58

30 Vibratory Moments, Configuration I, VI and UH-1H.... 61

31 Vibratory Moments, Configuration II and UH-1H. 62

32 Vibratory Moments, Configuration III and UH-1H. 63

33 Vibratory Moments, Configuration IV, V and UH-1H .. 64

34 Steady Moments - Typical for All Configurations. 65

35 Centrifugal Force vs. Blade Radius. 66

36 Weight Distribution . 67

37 Flatwise Stiffness Distribution.68

38 Edgewise Stiffness Distribution.69

39 Torsional Stiffness Distribution. 70

40 Blade Static Deflection Comparisons. 71

41 Blade Flexural Axis Comparisons. 72

42 Blade Center of Gravity Comparisons. 73

Figure Page

43 Typical Goodman Diagram. 86

44 Stress-Cycle Curve. 88

45 Material Strain Allowables. 90

46 .50 Caliber Hit Sikorsky Main Blade.95

47 Blade Spar Structural Damage.96

48 Blade Tear Damage.116

49 Blade Tear Damage. 116

50 Blade Gash Damage. 117

51 Blade Dent Damage.117

52 Impact of Blade Acquisition Cost,

1972 Configurations. 127

53 Impact of Blade Acquisition Cost,

1980 Configurations.128

54 Blade Acquisition Cost, 1972 - 1980. 145

55 Forecast of Material and Labor Costs. 146

56 Maximum Torsional Deflection ... 155

57 Normalized Vibratory Stress. 156

58 Rotor Thrust - Blade Twist Curve. 157

59 Remove Skin and Prepare Overlap Area. 205

60 Trim Patch to Fit.205

61 Prime Patch and Skin. 205

62 Prime Patch and Skin.205

63 Fill Edge Separations.206

Figure Page

64 Apply Adhesive and Position Patch.206

65 Apply Adhesive and Position Patch.206

66 Apply Adhesive and Position Patch.206

67 Positioned Patch. ,... 207

68 Apply Adhesive.207

69 Apply Adhesive to Overlay and Position

Scrim Cloth.207

70 Positior Overlay.207

71 Install and Inflate Compression Blanket.208

72 Finished Patch.208

73 Damaged Blade.210

74 Repair Materials.210

75 Replacement Plug.210

76 Plug in Place.210

77 Apply Foam to Cavity.211

78 Expanded Foam.211

79 Foam Trimmed and Sanded to Contour.211

80 Apply Adhesive to Overlay and Plug.211

81 Apply Overlay - Apply Adhesive to Plug.212

82 Apply Overlay to Plug.212

83 Apply Compression Blanket .212

84 Finished Repair.212

85 UH-1H Mission Environment.233

Figure Page

86 Impact of MTBR Criteria on Cost Effectiveness -

Acquisition Cost Sensitivity.250

87 Impact of Blade Acquisition Cost - Baseline -

1063 MTBR. 251

88 Plan for Future Hardware Evaluation.256

89 Outboard Specimen.259

90 Inboard Specimen.260

91 Stress Level of Operation.262

92 S-N Strength Data.262

93 Crack Propagation...262

94 Static Rap Test for Blade Frequency.263

95 2000 HP Main Rotor Test Stand.265

96 Rotor Hover Performance Comparison

Configuration V vs. Standard Blade.266

97 Typical Strain Gaged Blade.269

LIST OF TABLES

Table Page

I Comparison of Physical Properties. 74

II Blade Design Features. 75

III Material Properties. 77

IV Blade Stress in Level Flight Cruise -

Configuration UH-1H . 81

V Blade Stress in Level Flight Cruise -

Configuration I and VI. 82

VI Blade Stress in Level Flight Cruise -

Configuration II . 83

VII Blade Stress in Level Flight Cruise -

Configuration III. 84

VIII Blade Stress in Level Flight Cruise -

Configuration IV and V. 85

IX Structural Analysis for Various Modes of

Failure /Damage - Configuration I . 92

X Structural Analysis for Various Modes of

Failure/Damage - Configuration IV. 93

XI Reasons for UH-1D Blade Removal -

MTR/MTBR Analysis. 103

XII Reliability Apportionment - Baseline

UH-1D Blade . 108

XIII Failure Rate Summary . 112

XIV Blade Repairability and Level of Maintenance. 114

XV Repairability Summary. 114

XVI Aircraft Cost Effectiveness - 1972 . 124

XVII Fleet Effective Cost - 1972 . 124

xiii

Table Page

XVIII Aircraft Cost Effectiveness - 1980 . 126

XIX Fleet Effective Cost - 1980. 126

XX 1972 Cost Effectiveness Summary - Baseline. 130

XXI 1980 Cost Effectiveness Summary - Baseline. 131

XXII 1972 Cost Effectiveness Summary - Conf. I. 132

XXIII 1980 Cost Effectiveness Summary - Conf. I . 133

XXIV 1972 Cost Effectiveness Summary - Conf. II. 134

XXV 1980 Cost Effectiveness Summary - Conf. II. 135

XXVI 1972 Cost Effectiveness Summary - Conf. Ill. 136

XXVII 1980 Cost Effectiveness Summary - Conf. Ill. 137

XXVIII 1972 Cost Effectiveness Summary - Conf. IV. 138

XXIX 1980 Cost Effectiveness Summary - Conf. V. 139

XXX Cost Effectiveness Summary . 140

XXXI Cost of New Blade to the Army. 141

XXXII Reliability Analysis - Configuration I. 159

XXXIII Repairability Analysis - Configuration I. 163

XXXIV Math Model R/M Input Variables - Configuration I.... 167

XXXV Design Failure Mode and Effect Analysis -

Configuration I . 168

XXXVI Reliability Analysis - Configuration II. 172

XXXVII Repairability Analysis - Configuration II. i76

XXXVIII Math Model R/M Impact Variables -

Configuration II. 180

xiv

Table

Page

XXXIX Design Failure Mode and Effect Analysis -

Configuration II . 181

XXXX Reliability Analysis - Configuration IV. 185

XXXXI Repairability Analysis - Configuration IV. 189

XXXXII Math Model R/M Input Vairables -

Configuration IV.193

XXXXIII Design Failure Mode and Effect Analysis -

Configuration IV.194

XXXXIV UH-1 Rotor Blade Design Cost Comparisons.234

XXXXV Reliability Apportionment - Baseline UH-1

With 1063 Hour MTBR.238

XXXXVI Reliability Analysis - Configuration V

Compared to Baseline UH-1 Blade With 1063

Hour MTBR.240

XXXXVII Repairability Analysis - Configuration V

Compared to Baseline UH-1 Blade With
1063 Hour MTBR.244

XXXXVIII Math Model R/M Input Variables-Configuration V
Compared to Baseline UH-1 Blade With 1063 Hour
MTBR .248

XLIX Cost Effective Summary - 1063 MTBR

Baseline Configuration - 1980 .252

L Cost Effective Summary-1063 MTBR

Configuration V - 1980 . 253

LIST OF SYMBOLS

'att

'dep

B ds

B inv

B repl

B req

ret

B
B

B sa

BD„

rem

BDSdep

BLCC

BLCC UH

BO dep

blade set attrition, sets/FH
mission availability
installed blades per aircraft
blades lost to attrition
total damaged blades sent to depot
removed blades sent to direct support
initial blade spares
blade replenishment spares
blades requisitioned from inventory
blades retired from service
blades removed or installed
blades lost to scrappage and attrition

total blades damaged
blades externally damaged

blades inherently damaged

damaged blades removed from aircraft

damaged blades sent to depot from direct support

blade life-cycle cost, $

baseline TJH-1 blade life-cycle cost, $

damaged blades sent to depot from organizational level

xvi

dep

BR ds

BR off
BR f

BS
BS,

dep

BS c

°r

cont

"fly

damaged blades repaired, depot

damaged blades repaired, direct support

removed blades repaired off aircraft, organizational level

damaged blades repaired on aircraft

total damaged blades scrapped, all levels

damaged blades scrapped, depot

damaged blades scrapped, direct support

removed blades scrapped, organizational level

number of biades
coefficient of thrust
blade comaker cost, $

blade contribution to aircraft flyaway cost, $

C fuel

C inst

^isp

rem

req

fuel and oil cost per pound of fuel consumed, $/lb

cost of blade installation, organizational level, $

blade contribution to initial spares cost, $

average mission capability, ton-knots

cost of blade removal, organizational level, $

cost to requisition and obtain replacement blades,
organizational level, $

xvii

^rsp

^xx

°yy

acq

^dep

CG ds

CG 0
CGR

dep

CGR

off

CGR

CGS

dep

CGS

CGS,

dep

blade contribution to replenishment spares cost, $

distance between the point under consideration and the
chordwise blade neutral axis, in.

distance between the point under consideration and the
neutral axis perpendicular to the chordwise axis, in.

single blade acquisition cost, $

total blade contribution to replenishment GSE cost, all
levels, $

replenishment GSE cost, depot level, $

replenishment GSE cost, direct support level, $

replenishment GSE cost, organizational level, $

replenishment GSE cost per repair, depot level, $

replenishment GSE cost per repair, direct support level, $

replenishment GSE cost per off-aircraft repair,
organizational level, $

replenishment GSE cost per on-aircraft repair,
organizational level, $

GSE support cost per aircraft, depot level, $

GSE support cost per aircraft, direct support level, $

GSE support cost per aircraft, organizational level, $
cost of blade receiving and inspection, depot level, $

xviii

Clds

C^dep

CM ds

CM 0

CMR

off

CMR

CPOL

cost of blade inspection, direct support level, $

cost of off-aircraft inspection for blade disposition,
organizational level, $

cost of on-aircraft inspection for blade repairability,
organizational level, $

total blade contribution to maintenance cost, all levels, $

blade contribution to maintenance cost, depot level, $

blade contribution to maintenance cost, direct support
level, $

blade contribution to maintenance cost, organizational
level, $

mean material cost per blade repair, direct support
level, $

mean material cost per off-aircraft blade repair,
organizational level, $

mean material cost per on-aircraft blade repair,
organizational level, $

blade overhaul cost, depot level, $

cost of shipping preparation, depot level, $

cost of shipping preparation, direct support level, $

cost of shipping preparation, organizational level, $

aircraft fuel and oil cost, $

xix

CPOL

dep

off

CR.

dep

CSds

<*0

CSH

CSH,

cont

dep

‘fid
CSHF

CSHP,

CSHP

DH
DS

dep

baseline UH-1 life-cycle fuel and oil cost, $

cost of blade overhauls, depot level, $

cost of blade repairs, direct support level, $

cost of off-aircraft repairs, organizational level, $

cost of on-aircraft blade repairs, organizational level, $

cost to dispose of scrap, depot level, $

cost to dispose of scrap, direct support level, $

cost to dispose of scrap, organizational level, $

packaged blade shipping cost from field to CONUS, $

empty blade container shipping cost from field to CONUS, $

cost of shipping blades to depot, $

packaged blade shipping cost from CONUS to field, $

cost of shipping overhauled blades to field from depot, $

blade shipping preparation cost, depot level, $

blade shipping preparation cost, direct support level, $

blade shipping preparation cost, organizational level, $

down hours
depot support

E c

^ce

E m

E mUH
EA
EL

El.

F n

fe uh

FEC

f s

aircraft down hours per flight hour

modulus of elasticity of the component (material),
lb/in. 2

aircraft cost effectiveness, ton-knots/$
mission effectiveness, ten-knots
baseline UH-1 mission effectiveness, ton-knots
total axial stiffness, lb

total flatwise bending stiffness, lb/in. *
total edgewise bending stiffness, Ib/in. 2

allowable alternating stress, lb/in.

endurance limit of material at zero steady stress,
lb /in. 2

ultimate tensile strength of material, lb/in.
baseline UH-1 fleet effectiveness, ton-knots

fleet effective cost, $

average mission fuel flow, lb/FH
flight hours

stress, lb/in.

combined steady stress, lb/in.
combined vibratory stress, lb/in. 2

xxi

gross weight, lb

GSE

ground support equipment

horsepower

In.

inches

K inv

ratio of blade inventory spares to blade life-cycle
replenishment spares

KSI

1000 lb/in. 2

L a

aircraft service life, FH

blade scheduled retirement life, FH

pounds

LCC

aircraft life-cycle cost , $

lcc uh

baseline UH-1 life-cycle cost, $

LCC nv

UH-1 non-variable life-cycle cost, $

material cost,$

^es

steady edgewise moment, in-lb.

M ev

vibratory edgewise moment, in.-lb.

M fs

steady flatwise moment, in. -lb.

M fv

vibratory flatwise moment, in. -lb.

M inst

mean maintenance man-hours per blade installation, MMH

xxii

rt mg srtm

w 11» 'Jn-W '■t.r-'r-. -

Mfem mean maintenance man-hours per blade removal, MMF

M re q mean maintenance man-hours to requisition and obtain

a replacement blade, organizational level, MMH

MEK methyl ethylketone

MIdep mean maintenance man-hours per blade receiving and

inspection, depot level, MMH

MI ds mean maintenance man-hours per blade inspection, direct

support level, MMH

MI q ^ mean maintenance man-hours per off-aircraft blade

inspection, organizational level, MMH

MI Qn mean maintenance man-hours per on-aircraft damaged

blade inspection, organizational level, MMH

MMH maintenance man-hours

MR dg mean maintenance man-nours per blade repair, direct

support level, MMH

MR 0 fl: mean maintenance man-hours per off-aircraft blade repair,

organizational level, MMH

MR Qn mean maintenance man-hours per on-aircraft blade repair,

organizational level, MMH

MS margin of safety

MSdep mean maintenance man-hours per blade scrappage, depot

level, MMH

MSds mean maintenance man-hours per blade scrappage, direct

support level, MMH

xxiii

MS q mean maintenance man-hours per blade scrappage,

organizational level, MMH

MTB a aircraft mean time between loss of blades to attrition -

flight hours

MTBd aircraft mean time between inherent or external blade

damage - flight hours

MTB e blade mean time between external damage, flight hours

MTB| blade mean time between inherent damage, flight hours

MTB g aircraft mean time between blade scrappage - flight hours

MTB ga aircraft mean time between scrappage and attrition -

flight hours

MTB sar aircraft mean time between scrappage, attrition, or blade
retirement - flight hours

MTBR mean time between removal, flight hours

MTR mean time to removal, flight hours

N fleet size

Nj-j total number of blades

N n number of cycles required to initiate a fatigue crack at

n n stress level

N r revolutions per minute

n n number of cycles at a specific stress level

ORG organization

xxiv

PBR

dep

PBR

P3R

off

PBR

PBS

POL

R .
civ

mil

non

centrifugal force, lb

percent of damaged and removed blades sent to direct
support, %

percent of received blades repaired at depot level, %

percent of received blades repaired at direct support
level, %

percent of damaged and removed blades repaired at
organizational level, %

percent of damaged blades repaired on aircraft, %

percent of received blades scrapped at direct support
level, %

percent of damaged and removed blades scrapped at
organizational level, %

petroleum, oil and lubrication

minimum combined stresses/maximum combined stresses
civilian maintenance personnel labor rate, $/hr

recurring costs,$

mission reliability

military maintenance personnel labor rate, $/hr
non-recurring cost,$
rotor radius f ft

xxv

R aircraft mission abort failures per flight hour

r labor cost, $/hr

SH shop-hours for fabrication

SL sea level

SSE special support equipment

STD standard

T d average daily downtime , hr/day

T m average mission flight time, flight hours

U a aircraft annual utilization, flight hours

U d average daily utilization, FH/day

V knot (nautical miles/hr) or velocity (ft/sec)

<f solidity = % of total blade area/rotor diameter area

xxvi

wm* iwe wti n

INTRODUCTION

The need for redesigning rotor blades for a combat environment is evi¬
dent from field experience with the UH-1 helicopter. The data show
that 30% of all UH-1 blades are scrapped in the field, and 58% are re¬
turned to depot for overhaul. However, 40% of the blades are scrapped
at overhaul. This means that 70% of the original blades are replaced by
new blades. However, of new and repaired blades, only 5% or less ever
last 2,000 hours.

In recognition of this, the Eustis Directorate has funded Vertol Division,
The Boeing Company, to investigate sectionalized blade concepts which
could be disassembled. The damaged sections could be scrapped and re¬
placed. Kaman Aerospace Corporation was funded to investigate methods
of making blades more field repairable. Sikorsky was funded to assess
the practicality of a field replaceable bonded pocket for the CH-54B, and
Sikorsky and Kaman were funded to investigate expendable blade designs.

This report presents Sikorsky's study of expendable blade designs. An
expendable blade is defined as a blade with a low enough unit cost that it
is more cost effective to throw it away than to send it back to depot.

This simple concept of expendability was expanded. In addition to low
unit cost, the blade must be damage tolerant to reduce the scrappage
rate. Secondly, the blade should be reliable enough to minimize non¬
combat malfunctions. The blade should also be highly repairable in the
field. If these goals are met, the blade would have a lower unit cost than
conventional blades and would be continued to be repaired in the field un¬
til its useful life is expended and it is finally thrown away. To quantify
this, the cost effectiveness of the Bell blade was compared with cost
effectiveness of the new designs with and without depot repair. Blades
were considered expendable when the cost effectiveness was higher to
repair in the field or scrap than to send blades back to depot.

The study was conducted in two time frames, primarily because the cost
of labor and materials is estimated to change significantly. In addition,
new composite technology which has not been demonstrated could not be
considered for near-term applications. An important part of this study
is to separate those designs which can be flown early and put into pro¬
duction in the 1970's. There are concepts, particularly the composites,
which will have a greater potential but require demonstration of cost,
manufacturing, and structural viability.

The report will first describe the methods developed to evaluate expend¬
ability; designs of aluminum, steel, and composite blades; and selection
of expendable blades. The report also includes a recommended plan for
hardware development.

DEVELOPMENT OF METHODOLOGY AND DESIGN CONFIGURATIONS

PROGRAM APPROACH

Several approaches were considered feasible to obtain blades of higher
cost effectiveness than the UH-1H blade. One possibility was to simplify
the present UH-1H design by reducing the number of components while
still maintaining a similar low-cost aluminum extrusion. Savings would
result from having fewer parts to fabricate and fewer man-hours for
assembly of each blade.

Another possibility was to increase the repairability of the present blade
by replacing some of the aluminum with fiberglass components. Studies
made by Sikorsky Aircraft and by Kaman Aerospace Corporation, Ref¬
erence 1, have already shown that fiberglass is not only more repairable
than aluminum but that it can be repaired in the field. For example, one
utilization would be the substitution of fiberglass for the aluminum trail¬
ing edge spline and skin, an area where most blade damage occurs.

Another approach was to investigate an entirely new type of structure,
i.e., complete fiberglass blades with their many advantages of higher
fatigue life and even higher repairability than the metal blades because
of their greater area of repairability. The higher fatigue life of fiber -
glass also offered the potential of increased aircraft performance not
possible with the limited properties of steel and aluminum.

Automation is not only another possibility, it is a necessity if blade cost
is to be reduced. The fiberglass design, for example, would be virtually
eliminated from contention, pricewise, without some form of automation
because individual layup of fiberglass sheets requires considerably more
man-hours of assembly time than a metal blade. In this respect, ex¬
truded fiberglass shapes and tape laying machines must be strongly con¬
sidered to minimize fabrication labor. The sheet metal design being con¬
sidered for study must also be automated to obtain low-cost parts. This
may be accomplished by a progressive die sheet metal rolling machine
with automatic feed and cutoff.

The approaches above were considered for both the 1972 and 1980 time
frames, taking into account the changes in material and labor costs for
the two periods.

RELIABILITY AND MAINTAINABILITY METHODOLOGY

The methodology employed to conduct the reliability and maintainability
portion of the study consisted of establishing an accurate and statistically
valid data base from which a complete reliability and maintainability
analysis of the baseline UH-1 main rotor blade was conducted. This
analysis then served as the basis for generating reliability and maintain¬
ability cost-effectiveness values for use in a mathematical model which
combined them with other design factors to determine a baseline cost-
effectiveness/life-cycle cost value. Candidate expendable blade designs
were then analyzed using the same procedure and compared to the esta¬
blished UH-1 baseline to determine changes in blade life -cycle cost
and aircraft cost effectiveness.

Reliability and Maintainability Data Research

All available reliability and maintainability data were collected and re¬
viewed. Collected data were screened to determine applicability to the
study in terms of equipment operational environment and similarity to
UH-1 blade design. Data originating from sources not representative of
UH-1 operational environment or blade design were discarded.

Reliability Data

The primary source of reliability data used as the basis for this study
was Reference 2. Initial research and screening of available data
sources revealed this document to be the most authori tative and valid
source of reliability data relative to the UH-1 main rotor blade in its
operational environment.

Maintainability Data

Background UH-1 maintainability data were collected from a variety of
sources. Repair limitations for the UH-1 main rotor head were esta¬
blished using References 3 and 4. Maintenance task times were cal¬
culated on the basis of the procedures set forth in these publications
relative to the UH-1 main rotor blade. Overhaul and restoration task
times were developed through Sikorsky overhaul and repair facilities
and reference to USAAMRDL furnished values and publications.

Baseline UH-1 Blade Profile

A baseline UH-1 blade reliability and maintainability profile was de¬
veloped using the assembled data. Specific failure modes were associat¬
ed with blade component parts, part repairability, repair levels, required

support equipment and related cost factors. These parameters were
then used to determine applicable cost variables for inclusion in the math
model to establish baseline UH-1 blade life-cycle cost and impact on
aircraft cost effectiveness. The math model also developed a list of
design sensitivities to provide direction in the design of cost-effective
candidate expendable blades.

UH-1H BLADE COST ANALYSIS

The task of producing a more cost-effective rotor blade for the UH-1H
aircraft is a challenging one, since the present UH-1H blade is already
quite low in initial cost. Low-cost aluminum and steel components are
already utilized in the construction of the present UH-1H blade, so
changes in material alone will not provide significant cost reduction.

The present UH-1H blade is being fabricated in production for approxi¬
mately $15.00 per pound based on cost and weight of $3000.00 and 200
pounds, respectively. The components are mostly aluminum; they con¬
sist of a primary spar extrusion, side doubler plates, root end grip
plates, skin, and honeycomb core. Additional parts are steel doubler
plates, steel and cobalt abrasion strips and structural adhesive. Since
the average cost of the material, including the price of the honeycomb
and structural adhesive, is approximately $2.50 per pound, the remain¬
ing $12.50 per pound for the production blade cost must be the result of
recurring costs of processing the material; i.e., man-hours associated
with component machining, finishing, forming extrusions, subassembly
and assembly operations. Cost must also include amortization of non¬
recurring cost for design, ground and flight test and manufacturing tool¬
ing. In addition, of course, a profit must be realized.

Since the base material for the UH-1H represented a small fraction of
total cost, other areas of the blade had to be examined to determine
where the greatest costs were incurred. Investigation showed that
changes in nonrecurring cost had only a small effect on final blade cost
when considering production runs of 10,000 blades, whereas a much
greater impact was obtained by changes in recurring costs. The exam¬
ples below illustrate how changes in nonrecurring and recurring cost
affect blade cost. Equation (1) shows the basic parameters associated
with blade cost (shown without profit for simplification). Equation (2)
incorporates typical values for the UH-1H blade where R non = $10 ,

N = 10,000, r = $25/hr, M = $500, and SH = 96 hr. Equation (3) depicts
the change in blade cost by reducing nonrecurring expenses by 25% keep¬
ing the other values of Equation (2) constant. Equation (4) shows a 25%
reduction in shop hours for fabrication while retaining other values of
Equation (2) constant.

’^WS'jjfrww^f fww* wr**imrt^^. ,*.

Blade Cost = R non + R c = R non + M + (SH)(r) (1)

where R non = Nonrecurring cost, $

R c = Recurring costs = M + (SH)(r), $

N b = Total number of blades
r = Labor cost, $/hr
M = Material cost, $

SH = Shop-hours for fabrication

Blade Cost = $10 6 + $500+ (96 hr) ($25/hr) = $3000 (2)

(Base blade) 10,000

Blade Cost = . 75<$10 6 ) + $500 + (96 hr) ($25/hr) = $2, 975 (3)

< R noiT 25 % TCjOT -
reduction)

Blade Cost = $10 6 + $500 + . 75 (96 hr) ($25/hr) = $2,400 (4)

(SH-25% 10,000

reduction)

As shown by Equation (3), for a production run of 10, 000 blades, a 25%
reduction in nonrecurring costs from $1,000,000 to $750,000 reduces the
blade cost by approximately $25. 00, which is an insignificant amount.

On the other hand, assuming a labor cost at approximately $25. 00 per
hour (includes base rate, overhead, etc.), a 25% reduction in shop hours
from 96 to 72 hours reduces blade cost by approximately $600.00 as shown
by Equation (4). This represents a recurring to nonrecurring cost factor
of 24 to 1, which is significant.

A similar example would show that recurring material cost has higher
fluctuation then nonrecurring costs but is small compared to shop hour
changes.

Blade Cost = $10 6 + . 75($500) + (96 hr) ($25/hr) = $2,875 (5)

(M-25% 10,000

reduction)

Same 25% reduction to reduce material cost (Equation (5) ) results in a
blade cost of $2,875 or a savings of $125 over the base blade of Equation
(2). This $125 is five times greater than that obtained by reducing the
nonrecurring cost of Equation (3), but is approximately one-fifth the shop
hour cost savings of Equation (4).

Figure 1 shows how Equations (3), (4), and (5) were expanded to show the
cost changes (up to 200%) for each parameter. Each curve is a function
of one of the three parameters R non SH or M while the other two para¬
meters remain constant. The specific points of Equations (3), (4) and
(5), for a 25% reduction in the base values of Rnon» SH and M, intercept
each curve at . 75 along the abscissa.

It is obvious from the slopes of the plots that nonrecurring costs are the
least sensitive to change; material costs fluctuate slightly more than
recurring costs, while shop hours provide the greatest impact. These
slopes can be utilized to determine the changes required to the para¬
meters to control cost. For example, if a base blade costs $3000 and
the price of the material is increased by 100%, i.e. , if it doubles from
$500 to $1,000, the cost of the blade would increase to $3,500 if all other
parameters remain unchanged (this point is shown at the intersection of
the M plot and 2. 0 on the abscissa). To retain the same base blade cost
of $3000, either the noncrecurring cost or shop hours must be reduced
by $500. Since the slope of R„„ n is virtually a horizontal line, it cannot
be used to reduce costs by $500. The SH slope, however, can be used;
Figure 1 shows that $500 is equivalent to 21% on the SH plot. This shows
a 100% M cost increase is offset by a 21% SH decrease.

Summary Conclusions

Several summary conclusions can be made from the above investigation:

1. Amortized nonrecurring costs are extremely low for large pro¬
duction runs and have little effect on final blade cost. Doubling the
nonrecurring costs from $10^ to $2 x 10^ adds a mere $100 to the
blade cost.

2. Material cost for the UH-1H is inexpensive because of the utiliz¬
ation of low-cost steel and aluminum. The material is only slight¬
ly sensitive to change, i.e. , doubling material cost would increase
blade cost by less than 20%. This would indicate that some quan¬
tities of higher cost material can be utilized to obtain properties

of stiffness, strength, low weight, etc. , not possible with the
present materials. For example, if high cost carbon is judiciously
applied in small quantities for torsional and edgewise stiffness

Constant Values
N b = 10,000
r = $25/hr

sjbiioq - 3S03 apntg

96 hr 5106 $500

Figure 1. Impact of Nonrecurring, Shop Hours, Material Cost on Blade Costs.

requirements, it should not have a great effect on cost pro¬
vided carbon decreases in price as forecasted.

3. Shop hour fabrication cost is the most sensitive blade para¬
meter; a change in shop hours magnifies change in blade cost.
The most efficient way to reduce blade cost is to minimize
number of shop hours. On this basis, automation is the key.
Automation also offsets the use of higher cost material as noted
in Paragraph 2 above.

COST-EFFECTIVENESS METHODOLOGY

The value of rotor blade design is measured by the relationship of the
benefits it contributes to and the costs it imposes on the UH-1 aircraft in
Army service. With the exception of a few attributes such as safety and
human factors, most system characteristics can be quantified and inte¬
grated into an aircraft cost-effectiveness criterion:

Mission effectiveness

Cost effectiveness = Life-cycle cost (ton-knots/mega $) (6)

A cost-effectiveness model is used to relate rotor blade design character¬
istics and aircraft cost effectiveness. The general logic of the model is
shown in Figure 2. The maintenance burden, including spares require¬
ments, imposed by a rotor blade design on the organizational, direct
support and depot levels is the prime blade contribution to aircraft life -
cycle cost. This burden, in terms of blade logistics over the aircraft
life cycle, is described in Figure 3. A computer model incorporating
these analyses was designed and is used to obtain cost effectiveness of the
UH-1 equipped with the candidate blade designs. A detailed description of
the cost-effectiveness model with equations is presented in Appendix III.

A mission analysis program translates UH-1 weight/performance
characteristics into mission capability and fuel flow. This program is
also described in Appendix III. A MTBR of 914 hours was used for the
baseline UH-1 during the study. A comparison of the most cost effective
configuration using a MTBR of 1063 hours instead of 914 hours is included
in Appendix V. The rationale for the use of ton-knots instead of ton-miles
is also included in Appendix V.

Expendability

The evaluation ol rotor blade designs depends basically on the aircraft
cost-effectiveness criterion. Absolute blade expendability implies that
scrappage and replenishment of a damaged blade is always more cost
effective than the minimum field level repair. This concept quickly be¬
comes impractical since even the cost of shipping a replenishment blade

Figure 2. Cost-Effective Model

**t nr« mmi

Blade

Mean Time
To Repair

Blade Abort
Failure Rate

Blade

Weight

Blade

Performance

Characteristics

Blade Non-
Recurring
Prod. Goat

Blade

Production
Run (10000)

Blade
Recurring
Prod. Cbet

Gen. Admin.
It Profit
Factor

1 -

Organizational Maintenance
Goat Factora

On - A/C Inapection
On - A/C Repair
Blade Removal
Off - A/C Inapection
Off - A/C Repair
Blade Requisition
Blade Inetallatlon
tXapoeltlon of Scrap
Shipping Preparation

Direct Sigiport Maintenance ]
Coat Factors

Blade Inspection
Blade Repair
Disposition of Scrap
Shipping Preparation

Depot Maintenance
l Coat Factora

' Shipping to Depot

Receiving li Inapection
Blade Overhaul
Disposition of Scrap
Shipping Preparation
Shippings to Field

Blade Repair

1 GSE Tooling [■ -

1 Container
Coat

Average
Ml a a Ion
Fuel Flow

Blade

Aoquialtlon

Coat

Shipping Coat -
Blade V Oont to
Field

Shipping Coat -
Container
From Field

Blade

1 Organizational
Maintenance Coat

Blade

Direct Support
Maintenance Co6t

Blade

Depot

Maintenance Cost

AIRCRAFT

Figure 2. Cost-Effective Model

,^‘ST^’T ??f <**! v " > ►7'^ v, r?^^ ! w*WiWWWrWM^y «wi»w««i»w«wm*» rrrTunyiiwmf^mpri <, _ „ FKm MJJBk

Crew

Cost

J Replenishment
[ GSE Cost

OomX

Effectiveness

ORGANIZATIONAL

DIRECT SUPPORT

Removed Blades
Repaired
Off Aircraft

Removed Blades
Scrapped

BUdes

Inherently

Damaged

BUdes

Externally

Damaged

Total Blades
Damaged

Damaged Blades
Repaired
On Aircraft

4 Damaged BUdes
Removed For
Maintenance

Removed Blades
Sent to Direct
Si 4 >port

ci '

Removed Blades
Not Repaired

Removed BUdes
Sea to
Depot

Blades

ReqUsltlooed
From Stores

BUdes Retired
On Scheduled

BUdes
Removed k
Installed

Received

BUdes

Scrapped

BUdes Sent
To Depot

Figure 3. Life-Cycle Blade Logistics.

11 -A-

can exceed a minor repair cost. A better measure of blade expendability
is the degree to which cost effectiveness is enhanced/compromised by
the elimination of depot level repair. The cost-effectiveness model is
used to analyze the candidate blade designs with and without depot level
repair. In the latter case, damaged blades normally sent to depot for
overhaul/repair are scrapped at the field level.

Fleet Effective Cost

Fleet effective cost is an equivalent measure of aircraft cost effective¬
ness on a fleet level. Defining the UH-1 with standard blades as the base¬
line aircraft, and the UH-1 with any candidate rotor blade design as a
candidate aircraft, then for a fleet of N aircraft,

{

N x Baseline effectiveness

Baseline cost effectiveness = N x Baseline life-cycle cost

Baseline fleet effectiveness
Baseline cost effectiveness = Baseline fleet life-cycle cost

similarly, Candidate fleet effectiveness

Candidate cost effectiveness = Candidate fleet life-cycle cost

(7)

( 8 )

(9)

The change in cost effectiveness can be generated by a change in fleet
effectiveness, fleet life-cycle cost, or most frequently, in both. Fleet
cost can be made an equivalent cost-effectivenesss measure by adjusting
fleet size to maintain baseline fleet effectiveness:

and

N ' - Baseline fleet effectiveness

Candidate aircraft effectivness (10)

Fleet effective cost = N' x Candidate aircraft life-cycle cost (11)
or

Fleet effective cost = Baseline fleet effectiveness _

Candidate aircraft cost effectiveness (12)

Mission Effectiveness

Mission effectiveness is theoretical mission capability degraded by mis¬
sion abort failures and unavailability. It is the product of mission
availability, mission reliability, and mission capability.

Mission Availability

Mission availability is the probability that the aircraft will be available
for a mission on demand. Rotor blade designs that increase maintenance
burden or mean time to repair will increase aircraft down hours per
flight hour and decrease mission availability.

Mission Reliability

Mission reliability is the probability that an available aircraft will be
able to avoid a mission abort due to system failure.

Mission Capability

Mission capability is a measure of how well an aircraft can perform its
intended missions. It is the mission effectiveness of a perfectly avail¬
able, perfectly reliable aircraft. For transport aircraft such as the
UH-1, mission capability is assumed to be the product of mission payload
and mission block speed (productivity) expressed in ton-knots. A change
in blade weight will cause a corresponding change in aircraft weight
empty. For some missions, this will change mission payload. Changes
in blade performance will change aircraft performance characteristics
and may affect mission payload, mission speed, or both.

For any given mission, a change in aircraft characteristics may not
always produce a change in mission capability. For example, the pay-
load demand for the selected mission may not be limited by the takeoff
payload capability of the aircraft. A decrease in weight empty would
yield no significant benefit for such a mission. On the other hand, if
gross weight limits payload or range capability, reduced weight empty
offers a substantial mission payoff. To obtain a realistic evaluation of
mission capability, a mission analysis program which simulates 1000
missions while varying payload demand, mission range, operating alti¬
tudes and temperatures according to expected probability distributions
was used. This program is described in Appendix III.

Life-P/cle Cost

The life-cycle cost of an aircraft is the total cost generated throughout
its service life. It is the summation of unit development cost, acquisi¬
tion cost, and operating cost. Essentially, it is a measure of total user
cost.

Unit Development Cost

Unit development cost is the total nonrecurring development cost of the

system amortized over the total number of aircraft procured. It is
assumed that the basic nonrecurring costs of the UH-1 aircraft have
already been fully amortized. Nonrecurring costs associated with dif¬
ferent blade designs are amortized over the specified 10, 000 blades to be
produced and are included, for convenience, in acquisition cost.

Acquisition Cost

Acquisition cost is the sum of vehicle flyaway cost, initial spares cost,
initial training and travel cost, and initial ground support equipment cost.

Vehicle Flyaway Cost

Flyaway cost is the direct cost of the aircraft without spares and includes
the acquisition cost of two blades. Blade acquisition cost includes non¬
recurring investment for engineering design, engineering test, and
manufacturing tooling amortized over 10,000 blades and recurring costs
for manufacturing labor, materials, and recurring tooling. Material and
labor requirements were estimated for the existing blade, and labor rates
and overhead factors were adjusted slightly to yield the contractually
specified $3000 blade acquisition cost. These labor rates and overhead
factors were used as a basis for evaluating the acquisition costs of the
candidate blades.

Initial Spares Cost

Initial spares cost includes the cost of spares in inventory and in the
supply pipeline. Blade initial spares are a function of blade replenish¬
ment rate and supply pipeline efficiency. It is assumed that blade initial
spares are proportional to blade life-cycle replenishment spares. The
proportionality factor was adjusted to yield the contractually specified
30% initial spares.

Initial Training and Travel Cost

This cost applies to direct personnel support of the aircraft and is as¬
sumed not to vary with blade design.

Initial GSE Cost

This cost covers the initial acquisition cost of ground support equipment
and is assumed not to vary with blade design.

Operating Cost

Operating cost is the sum of replenishment spares cost, fuel and oil cost,

maintenance cost, flight crew cost, and replenishment GSE cost. It is
the direct cost of operating the aircraft after acquisition for an entire
life cycle.

Replenishment Spares Cost

Replenishment spares cost includes the cost of blade replacements de¬
manded by scrappage and scheduled retirement. Blade damage rates,
repairability, and retirement life directly impact on replenishment cost.

Fuel and Oil Cost

Fuel and oil cost over the service life of the aircraft is a function of
average fuel flow. The impact of blade design on fuel flow is provided
by the mission analysis program.

Maintenance Cost

Maintenance cost includes the cost of labor, materials, and shipping on
the organizational, direct support, and depot levels. Blade design
influences the repair, scrappage, and shipping rates in the blade logis¬
tics analysis. In addition, the cost factors such as material cost per
repair, overhaul cost, and man-hour burden to repair are affected by
blade design.

Flight Crew Cost

This cost is assumed not to vary with blade design.

Replenishment GSE Cost

This cost includes consumable tooling used in blade repair and is affected
by the blade repairability scheme.

UH-1 BASELINE COST EFFECTIVENESS

Mission Capability

The mission analysis program described in Appendix III was used to
establish the following:

Mission capability 50. 344 ton-knots

Average mission fuel flow 533. 441° lb/hr
Average mission time 0. 51945 FH

Mission Availability

0. 7500

Based on a total of 4. 3796 down hours per flight hour.

Mission Reliability 0. 99224

Based on 0. 015 mission abort failures per flight hour.

Mission Effectiveness 37. 466 ton-knots

The product of the above mission parameters.

Ten-Year Life-Cycle Cost Per Aircraft

Baseline UH-1 life-cycle costs were estimated parametrically or
assumed as follows:

Unit development cost

$ 0

Flyaway cost

$300,000

Initial spares cost

$100,000

Initial training and travel

$210,000

Initial GSE cost

$37,000

Acquisition cost

$647,000

Replenishment spares cost

$150,000

Fuel and oil cost

$53,000

Maintenance cost

$255,000

Flight crew cost

$480,000

Replenishment GSE cost

$ 0

Operating cost

$938,000

Life-cycle cost

$1,585,000

Blade Life-Cycle Cost

Baseline blade life-cycle cost consists of the contributions made by blade
characteristics to the aircraft life-cycle cost:

Blade contribution to flyaway cost $6,000

Blade contribution to initial spares cost $1,998

Blade contribution to replenishment spares cost $36,202
Blade contribution to maintenance cost $4,259

Blade contribution to replenishment GSE cost 0

Baseline blade 10-year life-cycle cost per aircraft $48,459

Fuel and oil cost for the 10-year aircraft life can be computed on a sys¬
tem level from average fuel flow

Baseline fuel and oil cost = $53,344

The nonvariable part of UH-110-year life-cycle cost becomes

LCC nv = 1,585, 00 - 48, 459 - 53, 344 = $1, 483,197 (13)

The computation of 10-year life-cycle cost per aircraft for any candidate
blade design can now be established:

Candidate life-cycle cost = $1,483,197

+ Candidate blade life-cycle cost
+ System fuel and oil cost (14)

This specific approach is used in the cost-effectivness model.

MATERIAL SELECTION

An assessment of the various materials available for rotor blade con¬
struction was undertaken at the start of the design study. The candidate
materials were compared with respect to the cost per pound and are
shown in Figure 4. The more conventional materials, such as aluminum,
steel and "E" type fiberglass, are the obvious low-cost materials. Steel
and aluminum can be obtained for $1 per pound; the "E" glass varies in
price considerably dependent on its form; the raw spool roving and resin
are under $1 per pound while the cost of preimpregnated "E" glass
purchased in a woven cloth or lineated ply form may be several times
this amount.

Titanium sheet material would be about ten times higher in cost per pound
than the conventional metals, aluminum and steel. Titanium usage in an
expendable rotor blade can be justified only in very small quantities for
blades which require both a high strength and high strain allowable
material.

The "S" glass prepreg and woven fabric is five times more expensive per
pound then the "E" glass in the same form. Therefore, since "S" glass
properties are only 10 - 15 percent better than the considerably less
expensive "E" glass, the use of ”S" glass would not be cost effective.

The very high modulus to weight ratio materials such as carbon and boron
composites are costly at the present time. However, the projected cost
of carbon composites shows a very large potential for cost reduction in
the late-1970 time frame. Increased usage of carbon composites and

Base Material Cost (^lb)

Figure 4. Material Cost Comparisons.

improvements in production methods could result in material costs of
$25. 00 per pound. At this price level, a limited amount of carbon can
be cost effective in an expendable rotor blade.

DESIGN CONFIGURATIONS

Six basic design concepts are considered in this study: an extruded
aluminum spar (Configuration I), a rolled sheet metal blade of stainless
steel and aluminum (Configuration II), a composite blade with a "D"
shaped fiberglass and carbon spar with a fiberglass skin (Configuration
III), a composite twin beam spar blade with a fiberglass spar and a
carbon skin (Configuration IV), a composite twin beam spar design (Con¬
figuration V) similar to Configuration IV but having an automated
pultrusion trailing edge skin, and an aluminum extrusion spar (Configu¬
ration VI) similar to Configuration I but also having an automated
pultrusion skin. Each concept included many trade-off studies to opti¬
mize the design in such areas as spar chord, trailing edge construction
and materials. Duplication of the basic UH-1H properties such as weight,
stiffness and natural frequencies was the primary design criterion used
for each of the design configurations.

CONFIGURATION I - AS-EXTRUDED SPAR

The final version of Configuration I is shown in Figure 5. It consists of
a two-piece constant section aluminum spar with fiberglass composite
trailing edge skins and spline. The skins are stabilized by a resin
treated polyamide paper honeycomb (Nomex). The heavy leading edge of
the spar is one of the features of rnis design. Providing a iieavy leading
edge in the spar eliminates the need for separate counterweights to mass
balance the blade. The sidewalls of the spar were also increased to
eliminate the side doublers. The heavy leading edge of the spar also
provides increased erosion and foreign object damage tolerance because
more material is available to blend out nicks and other damage in the
leading edge. In addition, a polyurethane abrasion resistant coating is
placed over the leading edge to further reduce erosion. The trailing
edge fiberglass skins consist of "5" glass with two-thirds oriented at
± 45° and one-third oriented at 0' to provide maximum shear strength
in the skin for torsional rigidity and to resist the edgewise bending
shears. At the trailing edge, a buildup of 0° glass fibers is used in the
spline to provide edgewise bending stiffness. A plastic honeycomb was
selected for the trailing edge core because it is more damage tolerant
and less susceptible to corrosion than metal core.

Configuration I has a total of fifteen less components than the UI1-1H.

The two-piece spar replaces seven components consisting of two lead-

/— CC'sl^-rA »r ir^Ai oOt,l AlUMiIIUM alio.'

• Closure : itCt oOti aluminum alloy

Ufl-ntKhC lONAL E jLA

1 AS" ORIEHTATiN

VIEW/ E

jNiriAtCHO.'.Al. V UlASs 0° ORlL NTATIOI

SECT I CM E-b

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jP.- Pi Al h b r d Al .MiNUM

alloy

SECTION A A

t _^scau t t riLEL L>-ip E'late

n i 2 3 4

INCHES

Figure 5. Configuration I.

.LOr

/—ABRASION RESISTANCE
COATiNG (SHADED AREA)

1 UNI DIR EC 'ION l jLASS
TAb" ORIENTATION

- UNIDIRECTIONAL V A ASs
0 ° ORirNTAfiON

HONEYCOMB CORE
NOKO ^f,CFa- 2'^ T 3

DOUbLLR PL Alts
/-’L* GLASS tAb'

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INCHES

VI L AV F.

Gi-'P R -II

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w r^

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i S l \
‘NCHES

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INCHES

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V WEIGHTS

RECTlONAl'F* GLASb
IENTATION

scale

n TTirm
0 l z 34 S6?8
INCHES

mg edge counterweights, four spar doublers and the present spar. The
polyurethane abrasion resistant coating replaces four additional steel
and cobalt abrasion strips, and me root end is simplified by the removal
of six doubler plates.

This configuration is enhanced over the present UH-1H design by (a)
reducing the number of parts, (b) providing an as-extruded, machining-
free extrusion and (c) increasing the repairability of the trailing edge by
the incorporation of the fiberglass skins. The lower cost of Configu¬
ration I coupled with its higher repairability results in more cost-
effective blade than the UH-1 for both 1972 and 1980, as shown in Tables
XVI and XVIII.

This concept is a simplication of the present UH-1H blade, utilizing an
aluminum extrusion as the primary structural member. The first de¬
sign considered use of an open-ended "C" section with a very thick
leading edge whose outer surface formed the contour of the airfoil over
the forward 25 percent of the chord. This was combined with a struc-
ural trailing edge to form the blade as shown in Figure 6. This blade
spar concept facilitates manufacture because the counterweights and
side plate doublers are made integral parts of the spar, eliminating the
fabrication and handling of separate components.

It was possible to closely match all of the important parameters of the
UH-1H blade with the "C" spar except the torsional rigidity which was
approximately two-thirds the required amount. Since torsional rigidity
is very important in the performance and stability of a blade, it was im¬
proved by closing the open end of the "C" spar with an extruded aluminum
closure piece across the open legs of the spar (Figure 7). This increased
the torsional rigidity to the same level as the UH-lII blade.

Since the open end of the spar required closing the torsional rigidity, a
hollow (one-piece) extrusion was investigated as an alternate to the two
piece spar extrusion cited above. The study showed that closer toler¬
ances could be held on the open "C" section and it could be assembled as-
extruded with a minimum of machining. Tolerances ranged widely on
the hollow extrusion because of the difficulty of extruding a spar section
having a heavy nose and a thin backwall. The asymmetric hollow extru¬
sion tends to shift the mandrel out of position during the extrusion
process,causing poor dimensional control resulting in considerable
machining in the final process. The base cost of the hollow section ex¬
trusion was double the "C" section; also, because of the additional labor
required for machining the hollow extrusion, the total cost was tripled
over the ”C" section. On this basis, the ”C" section with the closure
piece was considered the most cost effective for the design. The trailing
edge portion of the blade is a structural fairing which completes the

" Spar

Figure 7. C" Spar Blade With Backwall Channel

chordwise airfoil contour aft of the spar. For high repairability and
ease of field replacement, segmented pockets approximately 1 foot wide
are preferred for the trailing edge. However, the major portion of the
blade’s inplane stiffness is provided by this trailing edge structure;
therefore, it must be continuous and strong as well as light in weight.
Various types of construction and materials were considered for support¬
ing and stabilizing the trailing edge skins. A study was made to deter¬
mine whether a honeycomb core or a rib-type construction would be the
most cost effective. The rib-type structure and honeycomb structure
were found to weigh about the same amount. However, the close rib
spacing required to develop the necessary skin panel strength resulted
in higher cost for producing the rib-type structure (Appendix I), since
many more parts must be produced and considerably more man-hours
are required for assembly. The study also showed that the honeycomb
is more repairable than metal or fiberglass ribs because honeycomb
can be repaired by simply injecting foam or bonding a new core section
in place (see Appendix II).

Foam was another consideration. Foam is desirable in small quantities
in areas where it is difficult to assemble or where the shape of the struc¬
ture is irregular and complicates the machining of the honeycomb. A
foam-in-place is ideal for this type of design. The foam also has the
advantage of being less expensive than the honeycomb, and its repaira¬
bility is excellent. However, to obtain the same stability and compres¬
sive strength as the honeycomb, a high density foam 25 to 30 pounds
heavier than the honeycomb presently used in the blade is required. This
increase in weight by itself would not be totally undesirable; it is the addi¬
tional 50 to 60 pounds of counterweight required in the leading edge to
mass balance the blade about the feathering axis which results in a total
blade weight increase of 75 to 90 pounds which is not acceptable.

The material for the skin and the trailing edge spline was chosen after
evaluating aluminum and fiberglass. Aluminum provided a trailing edge
structure that was 8 pounds lighter in total blade weight to produce the
same blade stiffness. Cost of the aluminum structure was also some¬
what lower than the fiberglass structure. However, when the overall
cost effectiveness of the two materials was considered, the fiberglass
proved to be more cost effective. Fiberglass provides a Higher degree
of repairability in the field and also operates at much higher margins of
safety than the aluminum. Because of these attributes, more trailing
edge damage is repairable in the field and the number of spares required
is reduced.

The fiberglass trailing edge structure could be fabricated several ways.
One method is by vacuum-injection molding; this is a method of laying

up cloth (without resin) in a half-mold which has the curvature of the
airfoil contour. After completion of cloth layup, a mating half-mold is
securely fastened and sealed into position. Resin is injected at one end
of the mold while a vacuum is drawn on the other end. The amount of
resin can be closely controlled to uniformly permeate the cloth to obtain
good repeatability of resin content. This method is more applicable for
fabricating whole sections or assemblies; for example, a complete
trailing edge section containing fiberglass ribs along with the skin or
layup of an entire blade. However, for either design, there is a require¬
ment for considerable hand layup and the process does not appear feasible
for automation; therefore, it is relatively costly.

Another version would be to automate the skin layup by designing ma¬
chinery for the application. The skin plies could be produced in a skin
mold utilizing automated tape laying equipment; in this process, a machine
with the appropriate tape on a roll travels over the required length of the
skin mold laying down plies of prepreg. Microswitches to control lateral
and longitudinal motion and automatic cutoff and shutdown of the tape and
equipment at the completion of tape layup would be designed into the
mechanism to obtain complete automation. This present state of the art
method was selected for this design.

Filament winding is another method. The filament on a spool is coated
with resin as it is wound onto a tooling cylinder whose perimeter and
length are equivalent to the dimensions of two side skins of the trailing
edge. The process is completely automated so that filament can be
wound back and forth, at ± 45°, on the cylinder. After completion of
filament winding, while still in the wet layup, the skin is cut in two
pieces and laid in skin molds for curing at a specified time and tempera¬
ture. This process is presently being used very successfully.

The above methods outlined for fabrication of the skin can be applied to
the fiberglass spline; however, the spline with its longitudinal "E" glass
and constant solid section is adaptable to the pultrusion method. In this
process, filaments on spools are dipped in a resin and are pulled through
a spline shaped die. Since the filaments are of an indefinite length, the
process is a continuous operation fabricating approximately fifteen parts
per hour and reducing shop cost to a minimum. The pultrusion process
fabricates a constant cross section; however, the spline can be varied in
cross section by machining the outboard end to the desired taper.

The study showed that the constant outboard section of the blade accounted
for approximately 80% of the total cost; therefore, primary design effort
was to reduce cost in this area. However, some investigations were also

made of the blade root end to determine if a more cost-effective design
could be obtained by simplifying parts or reducing machining and assem¬
bly chop hours.

Figure 8 shows one side of the present UH-1H root end; it consists of a
number of thin sheets and one thick forging plate arranged in a stacked
fashion to obtain a smooth transition in cross-sectional area, increasing
from the constant outboard section to the inboard end of the blade. Fig¬
ures 9 through 13 show various concepts ranging from a completely
laminated built-up section (Figure 9 ) to a one-piece aluminum die
forging (Figure 11 ). Another was a spar extrusion with an enlarged or
upset end (Figure 10 ) to provide the necessary increase in cross sec¬
tional area. Two approaches similar to the present design were: Fig¬
ure 12, which replaced the aluminum doublers with fiberglass layup while
retaining the same grip plate, and Figure 13, which eliminated some of
the doublers by extending the grip plate.

Figure 9 removed the need for machining by eliminating the forging, but
it was less cost effective because the additional pieces increased (a) the
shop hours for assembly, (b) the amount of adhesive required for bonding,
and (c) failure modes. The initial cost of the spar extrusion of Figure 10
was found to be costly,and the poor tolerance control on the extrusion re¬
quired additional machining expense. Figure 11 simplified assembly by
replacing all root end components with a one-piece forging on each side.
However, experience has shown that it is not good practice to bond thick
pieces together because of the need of extremely close tolerances on the
mating parts to provide a good bond joint. These close tolerances result
in costly fabrication and the possibility of a high rejection rate of out-of¬
tolerance parts. Figure 12 consists of a layup of fiberglass prepreg
molded with a flat surface as shown to facilitate machining and subsequent
bonding of the aluminum grip plate. The study showed that simplification
of the grip plate did not offset the additional cost of the fiberglass mater¬
ial or the increased assembly time of the fiberglass layup and was there¬
fore less cost effective. Figure 13 is the most similar to the UH-1H de¬
sign; however, it was simplified by eliminating a total of six doublers
while slightly extending the present drag plate. It was shown to be slight¬
ly more cost effective than the (JH-1H and was selected as the final root
end configuration.

Figure 10. Stepped Extrusion Root End.

Figure 11. Solid Aluminum Root End.

Figure 12. Fiberglass Laminates Root End.

Figure 13. Reduced Doubler Root End.

CONFIGURATION II - SHEET METAL ROLL-FORMED SPAR

Roll-formed stainless steel plate and aluminum sheet are utilized in this
concept to build up a bonded spar cross section. The low cost of alumi¬
num and steel sheet material coupled with the opportunity for a highly
automated production line are attractive features of this design concept.

Flat plate stock on a continuous roll and of the required thickness is
processed through a multiple stage rolling mill. This is a process, as
shown in Figure 14, where the material is passed through roll stands,
equipped with contoured roller dies, and formed by stages into an ulti¬
mate desired shape. The process produces parts which can be held to
very close tolerances. After forming, each component is cut to required
length with an automatic cutter coordinated to the stage rolls. These
cut-off machines can cut formed metal with minimum distortion and in
most cases into lengths close enough to specification that no further cut¬
ting or trimming is required. This is a high-speed process; the equip¬
ment operates at 50 feet per minute producing approximately 100 parts
per hour, resulting in extremely low fabrication cost.

After all sections have been formed, the parts are assembled by stacking
the three components together after automated tape laying equipment has
applied adhesive to the surfaces to be bonded, as shown in Figure 15.

Blade twist is built into the spar during the assembly operation; the open-
ended sections are easily twisted prior to bonding and are locked in the
twisted condition forming a closed tubular structure after adhesive curing.

The steel leading edge of the spar is made sufficiently thick to provide
the counterbalancing moment required to balance the outboard portion of
the blade without additional balance weights. This feature reduces the
number of blade component parts by making leading edge counterweights
unnecessary.

Since a trade-off study for Configuration I showed that a combination of
fiberglass skin, fiberglass spline, plastic honeycomb and the simplified
root end design of Figure 13 were the most cost effective, these concepts
were also used for Configuration II, as shown in Figure 16. The combi¬
nation of roll forming the spar and tape laying of the fiberglass skins re¬
sulted in almost a completely automated blade. The cost model study
showed this configuration to be slightly less cost effective than Configu¬
ration I but more cost effective than the UH-1H blade for both 1972 and
1980, as shown in Tables XVI and XVIII.

r°

.RIP PLATE GOfei R.UMI.NUM ALLOY

« DOUBicfi PLATES GObl ALUMINUM AaCY

UNiOiREC TION f' OLA|
IAS' OR it NEATlOII

*— ON AC PLATE fcOGi AiiNMUU ALLOY
INBOARD CAP
INBOARD SHNi WLUITS

HONtYC.CMS COLL
V0NLX ^dlL

DOUBLER PLATE
/-’E OtASS I

ABRAMTN PCS6TANT RCUMtTHMt CCWMS

SPAR C StC ION STAINLESS STEEL
- SPAS CHAT. Itt Ki» Al JMlIJUM ALLOf

SPA a Channel icw aluminum alloy

UNOiAECT/JEiAl E‘GLASS (f ORIENTATION

UNIDIRECTIONAL E GlASS
I AS" ORIENTATiON

SECTION B-b

SEE ENLMGtD V t h (

GRIP PI Alt GOblA UMIIOi
AHOY

LIPAj PLATE sOtjl ALUM.NUM ALLOY

ROOT END DOUBLERS

SECTION A-A

STEa GRIP a ATE

Figure 16. Configuration II.

j 5 ” susv**.

ABRASION RESISTANCE
COATING (SHADED AREA)

CHORDWlSE BALANCE
^WEIGHTS

ST* 2M ~

Tip CAP-

^WXvVv

JN'.EiREC’K'-N L wlAS j
LAS" ORIENTATION

LOUBLER PLATES

-’e" glass tab'

unidirectional e glass

0° ORIENTATION

SCALE

mTTTm
C I 2 34 ScTE
INCHES

VIEW E

L'E' GLASS 0“ ORIENTATION

-PiP PLATE

SPAR C SECTION

t S-| f

rail

“SPAR CHANNEL

SECTION C-C

0 12 3 4

INCHES

LUMINUM ALLOY

DRAG PLATE

UNIDIRECTIONAL L GLASS
0° ORIENTATION

UNIDIRECTIONAL E GLASS
iA5° ORIENTATION

SECTION D-D

0 12 3 4
INCHtS

TUw.IJTJZn

irectional'e' Glass
? It N TAT ION

scale

rnTTim
C I l 34
INCHES

i { I r c ? w

EP"Tj. J

SECTION C-C r . sc.au...,,

0 12 5 4

INCHIS

.ASS

/ UNIDIRECTIONAL e'glass
/ 145° ORIENTATION

■’.UTlJa*

SCALE

I—I—I—I—I
0 12 3 4

INCHES

|h^b|

•vbb

W • ’’ " * '*

n^ii

1MBSS3I

CONFIGURATION III - FIBERGLASS TUBULAR SPAR

A "D" shaped tubular type spar made of fiberglass and carbon was con¬
sidered for Configuration III as shown in Figure 17. The use of a carbon
composite in an expendable rotor blade would not seem justifiable in view
of the high material cost. However, to provide the required torsional
and edgewise stiffness and also to maintain the required weight, a high
modulus fiber composite must be used. Projected reductions in the cost
of carbon, coupled with improvements in the rotor blade performance
capability and the opportunity for a highly automated fabrication tech¬
nique, should result in a cost-effective solution.

Unidirectionally oriented "E" type fiberglass is the primary spar ma¬
terial since it provides the maximum spanwise stiffness and strength at
the minimum per-pound cost. Carbon fibers oriented at ± 45° to the
spanwise axis are wrapped around the fiberglass spar to provide the
maximum torsional rigidity. The airfoil contour is completed by a
fiberglass skin which encloses the spar and extends to the trailing edge
of the blade. A unidirectionally oriented fiberglass spline, located in¬
side of the skin at the trailing edge, is used to increase the inplane
stiffness of the blade. Nomex honeycomb core is used to support the
skin aft of the spar. A polyurethane coating is applied to the leading
edge portion of the skin to provide abrasion protection.

The spar configuration was chosen to provide the maximum torsional
rigidity for a minimum weight of material. A tubular type structure is
the most efficient member for transmitting torsional loads and was
therefore selected for study. The chordwise dimension of the spar was
determined by conducting a study to find the most efficient percentage of
the blade chord to make the spar. The study showed that both torsional
stiffness and the flatwise bending stiffness were a maximum for a given
amount of spar cross-sectional area when the spar was 55-60 percent of
the blade chord. This is illustrated in Figure 18 which is a plot of spar
chord as a percentage of blade chord vs the reduction in stiffness from
the peak values obtained at 55-60 percent chord. For our design we
used a spar chord of 50 percent, a value close to the peak but which still
approximately maintains the mass balance at 25 percent chord without
the use of any additional counterweights.

The unique feature of this design concept is the method of producing the
spar tube. Present state-of-the-art fabrication techniques for produc¬
ing a tubular type spar from composites involve a layup of material a-
round a mandrel. The required thickness of material is built up on the
mandrel one ply at a time, possibly by automated tape laying equipment.
The finished layup is then cured under heat and pressure to produce the
finished spar. When a large number of plies are used, the process can

be lengthy even by using tape laying automation.

We propose the pultrusion method of producing a hollow composite spar.
Because the spar is relatively thick in cross section and composed prin¬
cipally of fibers running axially (a configuration lending itself to the
pultrusion technique), the risks involved in pultruding the spar are mini¬
mal However, the control of contour tolerances possible with the pul¬
trusion process for a hollow section is probably not sufficiently accurate
at this time to produce a spar to finished dimensions. Also, twist must
be molded into the spar since it cannot be warped into the twisted shape
after complete curing. The solution to both of these problems is to final
form the spar after the extruding process. A closed heated die, made to
the final spar contour and including required spar twist, is used to final
form the spar. Since it is possible to soften most epoxy resins after
they have been partially cured by applying heat, the partially cured spar
pultrusion could be placed in the die for the small adjustments in shape
and twisting required. An inflatable bag would be placed inside the spar
to force it against the die surface while the die was heated and the spar
brought to a fully cured condition.

The fiberglass skins and trailing edge spline can be fabricated by the
methods outlined for Configuration I. Final assembly of the various
components is an adhesive bonding operation, very similar to present
blade assembly procedures. A half airfoil section mold is used as the
assembly fixture. Components are assembled in the mold with adhesive
film between each part. A vacuum is drawn on a pressure bag placed
over the blade assembly, and the bonding operation is performed in an
autoclave under applied heat and pressure. The root end is a laminated
structure similar to those described for Configurations I and II.

The present high cost of high modulus material for this configuration re¬
sulted in a less cost effective (Table XVI)blade than the UH-1 blade for
1972. Reduced carbon prices as forecasted for 1980 indicate that this
configuration will be more cost effective than the UH-1H for that time
frame, as shown by Table XVIII.

CONFIGURATION IV - FIBERGLASS TWIN BEAM SPAR

A fiberglass twin beam spar (Figure 19) with a full chord width outer
carbon skin to form the airfoil contour was investigated as Configuration
IV. The twin beam spar is composed of an upper and a lower spar beam
separated by a honeycomb core. The spar beams are constructed of "E"
type fiberglass oriented at 0°. The material is low in cost and has
excellent sti 'n allowable properties which enable the spar to withstand
much higher oiade deflections without damage than an equivalent metal
spar. This "reserve strength" capability of the fiberglass spar provides

* «*r~**i*?**a,-**

-grip Plate 606i aluminum alloy

i-unidirectional carbon

1 AS” OPIEN7A T I0N

4 DOUBLER PLATES 6061 ALUMINUM ALLOY

ZLJ-J'-]

/I pH, r^-'-

p //: 41

' JflA , ^ z • as

/// V A

1 // '■ Drag PLATl 6061 ALUMINUM ALLOY

f ■ INBOARD CAP
' -INBOARD SHIM WEIGHTS

AbRASlON RFS STANT POLYURETHANE COATMG

— HONEY r OMb CORE
NONEX^CELL^t*

UNID'RlCTIONAL
0° rs tNTATlON

( DOUH
-‘E

r spar, jnidikectional "E glass o'cp'Entation

yiEW E

unidirfciional carbon

145° ORItllTAI ION

UNIOWLCriOUAL 'c' GLASS
l45’ ORIENTATION 2 PLIES)
0‘ ORitNTATION (2 PLIES)

UMD.RtCTlONAL t GLASS
()" ORIENTATION

SECTION B-B

r-GRIP PLATE 6061 ALUMINUM
ALLOY

- SEE ENLARGED VIEW E

V///A

- ROOT END DOUBLERS

DRAG PLATE 6061 ALUMINUM ALLOY

SECTION A-A \

STEEL GRIP PLATE

0 i J 3 «

KWS

Figure 17. Configuration HI.

- BIDIRECTIONAL CARBON
1 45° ORIENTATION

ABRASION RESISTANCE
COATING (SHADED AREA)

^ CHORDWISE BALANCE
\ WEIGHTS

-H&NEYCOMfc CORE
NONFX >if. CELL ?l ^t>

UNIOiRtCTIONALE Gt ASS
0° CR lNTATION (2PlIS)

UN n HCCIIONAL VGLASS
145° ORIENTATION (2 PLIES)

□

D/ -UNlDlR'LCTiO'.AL
0° ORIENTATION

1 DOU’ LER plates

-’E GIASS 0°(2 RUES)

^1 I.|], ; ; I

145 (2 PLIES)

E' GLASS

scale

I I I I I T 'T i
0 l 2 54 S6 7B
INCHES

yiEW E

.its;

g\;0-ectional t Glass
0" ORIENTATION

W E

-Grip Plate

SPAR, UNIDIRECTIONAL E GLASS 0°ORIENTATION

UNIDIRECTIONAL CARBON 145°ORIENTATION

STA 28 /

s'/ s’ 7 £ // V / /-/ /Vvv ¥t it? / y /// -H^-^Pyr7~ry~ r r ■> tvt sir 7 -s z

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. S' /.

SECTION c-c

0.254

INCHES

IRAG PLATE 6061 ALUMINUM ALLOY

DRAG PLATE

UNIDIRECTIONAL '£' GLASS
C° ORIENTATION

UNIDIRECTIONAL ’E'GLASS
145° ORItNTATlON (2 FLIES)
0° ORIENTATION i2 pLIES'

SECTiON DO sule

I I 1 1 I

r ' Z i 4
INCHES

—unidirectional 'e' glass

0° ORIENTATION

scale

i i l m i 1 i i

OI23A567B

INCHES

lAL'E'GLASS 0°ORIENTATION

r UNIDIRECTIONAL CARBON 145°ORIENTATION

f , ■ . ' /I ' snnn nn * , vr>, rr^-rrrr

{ 4 - 'r ■ ■ ' {Ui't k ii»u<

m rrrrrrrrr / * ? ✓ ^

SECTION! C-C

SCALE

i— i—i i—i
0 i 2 3 4

INCHES

M E* GLASS
N

\UJ±ULLUIUT7Z12ZL

UNIDIRECTIONAL ’E‘GLASS
145° ORltNTATlON <2FLItS)
0° ORIENTATION 12 RlIES>

2ZZZ2Z2*

D-D

SlA l E

~!—I-1—I

12 3 4

INCHES

Tisgif

WTW3-

# ^CT"-O^

^®7T V*** .

FinUflLA^ VW«5Sri
_ L*tf/6*t<wr /a Q

Figure 18. Flatwise and Torsional Stiffness Change With Spar Chord Change.

an increased margin of safety in the event of projectile-induced spar
damage. The twin beam concept has the potential of being a fail-safe
spar design with redundant load paths, since either spar beam may be
severed and the remaining spar beam and skin will carry the centrifugal
and bending loads.

The simple design of the spar beam lends itself to a highly automated
type of fabrication. The constant cross section of the spar from the tip
end to the root end of the blade makes two different construction methods
feasible. Automated tape laying machinery can be used to lay the fiber¬
glass preimpregnated tape into a mold. The layup is then cured in an
autoclave producing finished spar beams molded to the required contours.
An alternate method of spar fabrication which is even more attractive
than the molded spar is the pultrusion process previously mentioned to
fabricate the fiberglass spar and spline. The predominantly axial orien¬
tation of the spar fibers is particularly well suited to the automated con¬
tinuous nature of the pultrusion type process and should produce a spar
beam of high quality and low cost.

The skin must provide the torsional rigidity for the blade, since both the
fiber orientation and configuration of the spar do not provide very much
torsional stiffness. Carbon fibers in the skin are oriented at ± 45°, the
orientation at which the maximum torsional rigidity will be obtained.
One-third of the skin is comprised of unidirectional "E" fiberglass to
improve laminate strength properties, based on materials testing at
Sikorsky Aircraft. The complete wraparound skin is fabricated by two
female molds provided with the blade twist and is split on the chord line
so that the upper and lower halves of the skin are molded separately. In
this operation, the preimpregnated carbon and fiberglass tapes are laid
in the mold by automated tape laying equipment and then cured in an
autoclave.

The trailing edge spline and the leading edge skin have carbon fibers
oriented at 0° and are both adaptable to the pultrusion process. Both
these components would be fabricated in half sections by splitting on the
chordline to facilitate subsequent assembly and machining operations.

The blade is assembled in two halves as shown in Figure 20 by placing
the previously cured skin, spar beam, leading edge doubler, trailing
edge spline and Nomex core into the same skin molds prepared with struc'
tural adhesive. Each half of the assembly is bagged and cured in an auto¬
clave and then machined flat on the chordline and finally both halves are
bonded together into one assembly. A splice cap is bonded over the
leading edge skin to provide a means for shear transfer across the joint.

MBS AS
COAT

ODRECTCNAt 't'GlASS
CTORENWnON

« DOUBLER KATES 6061 ALIAWJM ALLOT

HONEYCQN® CORE
NOMEX CELL 2L

ttiOARD 9 TM WEIGHTS

CTo

1 7

UMDECTICMAL CARBON O’ ORENTA T ION

•ABRASION RESISTANT POLYURETHANE COATING
• LEADING EDGE MOLDED (XXNTERWHGHT
s\ -U^)R83CNAL CARBON Cf ORENTATION

UNCPECTONAL E'GLASS O’ ORENT ATION

/— LNCPECTIONAL CARBON 145’ ORIENTATION
(OUTER SKN)

■ABRASION RESISTANT POLYURE THANE COATING (REF;
yr RBEHGIASS SPICE CAP
V - OUTER SKN (REF)

NSV: LEACNG EDGE DOUBLER (REE I
NV\\r MOLDED COUNTERWEIGHT (REF)
Wy KYNMPEE,

SECTION-B B

2E ENLARGED VEW E

SLIP PLATE 6061 ALUMNW ALLOY
WEU£ PLATE faOEl ALM'EA' AU£Y

ROOT EIC DOUELERS

STEEL GRP R ATE

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JJUlLli

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jillllii

SECTDN A-A scale

r-T-r' r • i

0 I 2 i 4

INCHES

Figure 19. Configuration IV.

UNCRECTIONAL CARBON O' ORENTATION -

LMMETONAL fGLASS 6 ORENTATION

INCRECTlONAL CARBON 145’ORIENTATION
(OUTER SKN)

VIEW E

-grip Plate

/-WEDGE PLATt

STA 26

S.A i _» v *

tltll

C77Z

SECT

HOOT ENT) DOUBLERS

ENLARGED VEW E

-GRP KATE 6061 ALIMNLW ALLOY
— WEDGE PLATE 6061 ALLWNLW ALLEY

DRAG PLATE 6061 ALLJMNLW ALLCY

-DRAG PLATE

L GRP PI A'LE

(\ SCALE

INCHES

STA 288

^ UM3fECT0NAL CATO
O' ORENWION

sent

nnirrn

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WEDGE PLATE

r UNIDIRECTIONAL *L GLASS

SECTION C-C n

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INCHES

i PLATE

SECTION D-D

scale

r 7 ! TT

INCHES

'V*

Insert Counterweight,
Foam, Honeycomb and
Trailing Edge Block

Step 2/

Figure 20. Fabrication of Beam Concept,

The root end buildup for transfer of blade loads to the rotor head is sim¬
ilar to the other configurations. In addition, aluminum wedge plates are
bonded to the inside of the spar beams to provide greater bearing and
shear tearout strength in the area around the attachment pins. The wedge
plates, which are bonded to the spar beams prior to the assembly of the
two halves of the blade,form an interlocking taper joint with the root end
of the spar beams, as shown in Figure 21. The interlocking tapers pro¬
vide an additional load transfer path from the spar to the root end attach¬
ment that is independent of the primary bonded attachment. The large
amount of bond area between the outside doubler plates and the constant
section portion of the blade results in low shear stresses in the adhesive.
The large cross sectional area of the doublers,coupled with the higher
modulus of the aluminum wedge plates,results in most of the spar loads
being transferred into these components in the attachment area. These
components and attachment fittings then provide the load transfer capa¬
bility to the rotor head.

The half-mold fabrication principle provides ease of inspection and
assembly of the individual components. The open section provides higher
quality control because all members are set together in the mold, readily
inspectable at a glance. The primary structures such as the spar, lead¬
ing edge doubler, and trailing edge spline require no internal examination
like hollow extrusions or tubes because they are simple, solid sections.
The components are placed side by side in the mold without regard to
thickness tolerance because each half-assembly is machined flush on the
chordline after all components have been arranged in position. Fabri¬
cating in this fashion results in excellent dimensional control of the blade
aerodynamic contour after machining and final bonding of the two half-
molds.

The composite blades of Configurations III and IV offer a significant ad¬
vantage over metal blades in that they are inherently corrosion free.
Except for the aluminum root end, the blade material is nonmetallic.

These blades are field repairable. Composites can be prepared for bon¬
ding simply by sanding and cleaning the surface with solvent. This can
be done in the field at the direct support level. Bonded metal blades are
more costly to repair because they must be removed to a high echelon of
maintenance. The metal blade requires cleaning, priming, and bonding
under clean, atmospherically controlled conditions.

Configuration IV was found to be less cost effective than the UIl-1 blade
and Configurations I and II for 1972 because of the high cost of the carbon
composite. For the 1980 investigation, this configuration was made cost
effective by the reduced cost of the carbon and by fabricating a blade half

Fiberglass Beam

Figure 21. Fiberglass Root End Attachment.

section by the pultrusion process. The blade concept is shown in Con¬
figuration V, Figure 22.

CONFIGURATION V, TWIN BEAM FIBERGLASS BY PULTRUSION PRO¬
CESS

Configuration V is similar in design to Configuration IV, having all the
same attributes; in addition, improvements are made to the trailing edge
by eliminating the expensive honeycomb in this portion of the blade and
inserting a truss type skin as shown in Figure 22. The truss members
consist of outer skin, diagonals and inner skin. The outer skin is still
carbon at ± 45° orientation. The diagonals and inner skin are composed
of fiberglass sheets arranged at ± 45°orientation to provide torsional
rigidity and stability to the skin. The inner skin extends the full chord
of the blade, enveloping and conforming to the unidirectional carbon fore
and aft and the unidirectional fiberglass spar. The components of one-
half of the blade thus become one integral part; high and low modulus
unidirectional and cross materials are combined in one manufacturing
process.

This process is an extension of the simple pultrusion method of fabricating
the simple, solid spar and spline sections of the other configurations; it
is unique because it combines hollow and solid sections, unidirectional
and cross ply and dissimilar materials. Such a complex use of this pro¬
cess has never been demonstrated; therefore, it represents more risk
than the more conventional skin and honeycomb layup construction. How¬
ever, the potential for mass production of a trailing edge structure with
an internal system of longitudinal stiffening members, thereby elimina¬
ting the cost and assembly associated with honeycomb and skin attach¬
ment, represents a sizable reduction in production costs.

Figure 23 shows a schematic of the pultrusion process. The part being
fabricated is a simple, hollow, rectangular tube consisting of unidirec¬
tional glass between inner and outer layers of a cross-ply material. The
shape is obtained by a fixed mandrel extending through a curing die.
Fiberglass mat or cross-ply material on a roll is dipped in a resin and
then uniformly wrapped around the mandrel by a tunneling process as
shown. Filament "E" glass roving, also dipped in resin, is evenly dis¬
tributed over the mat material. A second layer of cross ply or mat is
then placed uniformly over the longitudinal filaments. The entire assem¬
bly is then slowly drawn through the curing die by a series of rollers
bearing against the outer surface of the mat material. The part is essen¬
tially cured by the time it exits the die. The complex structure of Con¬
figuration V utilizes the same basic principle as above; it is fabricated by
a series of triangular mandrels, one for each hollow in the trailing edge,
and a die conforming to the blade half section containing the spar, the

JNlDIR

GRIP PLATE 6061 ALUMINUM ALLOY

HONEYC

NIONEX?;

A DOUBLER Plates 606l ALUM'NUM alloy

DOUBLEP PlAIES
- CARBON iAS 1 (2 PLIES)

DRAG PLATE 606i ALUMINUM ALtOr

INBOARD SHIM WEKjHTS

HONEYC

NOMEX

'E GLASS (? PLIES)

ABRASION RESISTANT POLYURETHAAE CCWNG

UNIDIRECTIONAL V GLASS O'ORlEN TATlON

UNIDIRECTIONAL CARBON 145° orientation

/ (OUTER SKIN)

/ a UNIDIRECTIONAL'E GLASS *45° ORIENTATION

leading edge molded counterweight

>— FOAM i

UNOiR

CARBO

SECTION B b

unidirectional carbon
0°ORIENTATION (LEAWG EDGE
DOUBLER)

^-GRIP PLATE 6061 ALUMINUM
ALLOY

SEE ENLARGED VIEW E

ROOT END DOUBLERS

SECTION A-A

STEt GRIP PLATE

INCHES

Figure 22. Configuration V.

t a

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UNIDIRECTIONAL f GLASS O' ORIENTATION

HONEYCOMB CORE
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COATING (SHADED AREA)

CHORDWISE BALANCE

STA 288 '

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UNIDIRECTIONAL CARBON
145° ORIENTATION

UNIDIRECTIONAL CARbON
0° ORIENTATION

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1:45“ ORIENTATION

SCALE

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INCHES

; 2 PLIES)
PLIES)

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LEADING EDGE DOUBLER (REF)
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CARBON ORIENTATION

SECTION C-C

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SCALE

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INCHES

r-DhAG plate

UNIDIRECTIONAL CARBON

SECTION D-D

scale
i— r~i—i—i

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INCHES

CHORDWISE BALANCE

VwE IGHTS

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UNIDIRECTIONAL carbon
0° ORIENTATION

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UNIDIRECTIONAL 'E' GLASS r^TTTTTI

i■45° ORIENTATION inches

-ABRASION RFELSTANT PObURETHANE COATING (REF)
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SECTION D-D

SCALE
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trailing edge truss and carbon leading edge doubler and trailing edge
spline. The carbon and "E” glass materials would be arranged to enter
the die to produce the section of Configuration V.

CONFIGURATION VI,ALUMINUM SPAR- PULTRUSION TRAILING EDGE

Configuration VI (Figure 24) is a combination of Configurations I and V.
The aluminum spar is combined with an automated trailing edge pultru-
sion. The one-piece construction eliminates separate skins, trailing
edge spline and Nomex core. There is a period of development required
for this process; however, it is felt it can be accomplished by 1975. To
provide structural integrity, the blade is also equipped with Sikorsky’s
monitoring device BIMr. The blade spar is sealed and pressurized from
its root to a point just inboard of the counterweight retaining block. A
device to indicate pressure loss visually, by showing red, is installed
near the root end, where it is visible from the ground. This method of
continual surveillance of the spar provides ar a glance a more effective
structural inspection than x-ray. Although not shown, this same device
is applicable to Confit; u ations I and II.

This design combines all the advantages of the aluminum and fiberglass
concepts, resulting in a simplified design and reduced cost.

6IM 0 INDICATOR AND MANIFOLD

STAjgi

GRIP PLATE 6061 ALUMINUM ALLOT

A DOUBLER PLATES 6061 ALUMINUM ALLOY

) J 1

/ f*tl . 'I

/ 1 -A

— Drag plate 606i aluminum alloy
j *- IN0CARD CAP AND BIM^SEAL
' IN 60APU SHIM WEIGHTS

, DOUBl ER PLATES

/ E GLASS 145° (2 PLICS)

UNIDlREC
L45° ORI

fxip* w

E GLASS (2 PLIES)

VIEW E V ' E ' &LA ^S C? PLIES)

ABRASION RESISTANT POLYURETHANE COATNG

- CLOSURE PIECE 6061 ALUMINUM ALLOY

\ CONSTANT SPAR 6061 *—UNIDIRECTIONAL E' GLASS ;45°ORIENTATION

ALUMINUM ALLOY 1

\ ' / \\

unidirectional e glass

o° ORIENTATION

A~ •- .

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SECTION B-B

GRIP PLATE 6061 ALUMINUM ALLOY

SEE ENLARGED VIEW E

kw«'/

ROOT END DOUBLERS

SECTION A-A

SCALE

i-r--T" - t - t

0 I 2 5 A

INCHES

STEEL GRIP PLATE

DRAG PLATE 6061 ALUMINUM ALLOY

JAAAAAA/ ■ A7C7Vv*SG.

Figure 24. Configuration VI.

'Mjnidirectional 'e'glass

145" ORIENTATION

SCALE

rm rrm
0 I 2 J4 Sfe7B
INCHES

(.2 PLIES)

=L ILS)

IRECTIONAL 'e'glass
MENTATION

>001 ALUMINUM ALLOY

CTIONAL E* GLASS
NTATION

LASS

scale

rrrr m in

0 I 2 54 5678
INCHES

ANALYSIS OF DESIGN CONFIGURATIONS

NATURAL FREQUENCIES

The natural frequency of each design configuration was determined for
comparison with a similar calculation for the UH-1H rotor blade. Dupli¬
cation of the UH-1H frequency characteristics is considered of prime
importance in the design of a blade which would be compatible with the
UH-1H aircraft. First and second mode flatwise and edgewise frequen¬
cies were determined for the appropriate pin ended and cantilevered
ended blade root end conditions for the entire range of rotor rotational
speeds. For a teetering rotor, the odd-numbered harmonics in the flat¬
wise direction are pin ended and the even numbered are cantilevered
ended. Edgewise end restraint is always cantilevered. Since blade tor¬
sional response is of prime importance (Appendix I), the first torsional
mode frequencies were also determined.

Blade beam bending frequencies are determined using a computer pro¬
gram which considers the blade as a series of uniform rigid beam seg¬
ments connected by hinges and springs. Segment length, mass and
spring stiffnesses are varied along the blade to represent the nonuni¬
form characteristics of the actual blade. A system of nonlinear differ¬
ential equations describes the forces, motion and acceleration experi¬
enced by the segmented blade. The end conditions of the blade are intro¬
duced as constraints at the root and tip of the blade. Solution of a deter¬
minate system comprised of the differential equations of motion yields
the blade bending natural frequencies and modal amplitudes.

The Southwell plots for each of the blade configurations and the UH-1H
are presented in Figures 25 to 29. The plots show the close correlation
of each design with the UH-1H blade. In particular, the first edgewise
mode frequency and the first torsional mode frequency are of prime im¬
portance. Separation of the first edgewise mode from the one-per-rev
rotor operating speed is essential to preventing any resonance and ampli¬
fication of in-flight one-per-rev edgewise loadings. Each proposed blade
design is at least as far removed from one-per-rev as the basic UH-1H
blade.

LOAD DETERMINATION

Evaluation of the structural reliability of each of the four design concepts
requires a determination of the bending moments experienced by each of
the candidate designs. Since there is very little difference between the
mass and stiffness distribution of each design and the basic UH-1H blade,
we would expect the blade bending moments to be very similar. As
expected, calculation of the blade bending moments for each design at a

Frequencies - CPM x 10

0 100 200 300 400

Rotor Speed - RPM

Figure 25, Natural Frequencies .Configuration UH-1H .

' ,i. ..JiU i iUWW

m 3 w. i r * , i n t wvww' T ' V[ ' .-.r -p

representative flight condition showed only small differences between
designs.

The loads and moments were calculated on a UNIVAC^ 1108 computer,
programmed by Sikorsky Aircraft. Two computer programs are avail¬
able for determining the loads for a teetering rotor, a normal modes
analysis and a modified Myklestad analysis. The Sikorsky normal modes
transient analysis was modified to make it applicable to teetering rotor
blades. In this computer program,blade flatwise, edgewise and torsional
elastic deformation are represented by a summation of normal mode
responses. The modal equations of motion are integrated numerically to
permit rational coupling between airloads and blade response. In order
to treat the teetering rotor system program,modifications were made to
accommodate the blade mode shapes and natural frequencies characteristic
of the two-bladed configuration. The modal excitations of two blades
(located 180° apart in azimuth) were combined to give the generalized ex¬
citation of the teetering rotor system. Three hinged and three hingeless
flatwise modes, one hinged and two hingeless edgewise modes and two
torsional modes were retained in the analysis.

Correlation studies of the blade bending moments determined using this
program with those obtained from flight test data for the UH-1H aircraft
showed poor correlation of the edgewise moments. Computed edgewise
moments were much greater than the flight test data and proved to be
very sensitive to changes in first mode edgewise frequency. Because of
this sensitivity to edgewise frequency and some question as to the actual
hub stiffnesses of the UH-1H rotor system (which would directly influence
edgewise frequency), we elected not to use the normal modes analysis.

The Myklestad analysis program proved to correlate better with the test
data than the normal modes program and was therefore used for the blade
bending moment calculations. Good correlation of this program had been
previously demonstrated for the UH-1A teetering rotor as reported in

Reference 5. The Myklestad analysis is divided into two distinct parts.

The first is an aerodynamics analysis to determine air loadings on the
blade for a given flight condition. The second part is a dynamic struc¬
tural analysis which calculates blade response to the air loads by means
of a modified Myklestad approach.

The aerodynamic analysis calculates the rotor blade pitch, coning, and
tip path plane angles for given flight condition requirements (airspeed,
rotor speed, lift, drag, steady hub moments, density altitude). For the
aerodynamic analysis, the blade is assumed to be rigid with no flapping.
Flexibility effects are handled in the structural analysis. Two-dimen¬
sional steady-state airfoil lift, drag, and pitching moment data are used.
These include stall and Mach number effects. Air loads are calculated

at small azimuth intervals for chosen blade radial stations. The azi-
mutial harmonics of the loads are calculated and used in an iteration
procedure to trim the rotor to the flight condition requirements. The
final converged harmonic distribution of air loads is used in the struc¬
tural analysis to determine blade response.

The structural analysis is based on a lumped mass Myklestad approach.

It is a fully coupled flatwise-edgewise-torsional analysis. Influence co¬
efficients between masses are calculated using blade area moments of
inertia and beam theory. The method of solution is to determine the mo -
me nts, forces, deflections, and slopes at the blade root due to unit dis¬
placements and slopes at the blade tip, and due to the applied air loads
(corrected by damping loads due to blade deflections) and centrifugal and
other effects. The moments, displacements and slopes required to sat¬
isfy the root end boundary are then calculated, and the forces, moments,
stresses, deflections and slopes at each radial station are calculated
using these values.

Blade bending moments were calculated for a typical UH-1H cruise con¬
dition in level flight. A velocity of 90 knots, gross weight of 8, 500 pounds
and a rotor speed of 318 RPM were used for the calculation. Figures 30
to 33 illustrate the vibratory flatwise and edgewise bending moments ob¬
tained for each of the design configurations and the UH-1U blade. Figure
34 shows the steady moments typical for all configurations. Centrifugal
force of each of the blades is shown in Figure 35. Only one curve is
shown for the steady bending moments because it is representative of all
the blade configurations.

STRUCTURAL ANALYSIS

The objectives of the structural analysis work on the four design configu¬
rations presented in this report are to identify any major structural
problem areas and to evaluate the expected fatigue life for the designs
under study. All of the work is based on a direct comparison of the de¬
signs under study with a similar calculation for the UH-1H rotor blade.

The spanwise distribution of mass and stiffness, used in the determin¬
ation of the blade loads and stresses, is presented in Figures 36 through
39. Comparisons of spanwise deflection and flexural and center of
gravity locations are shown in Figures 40 through 42. Additional com¬
parisons of blade physical properties and design features are shown in
Tables I and II.

The properties of the materials under consideration for the configura¬
tions are shown in Table III. These properties were obtained from Ref¬
erences 6, 7, 8 and 9 and various material property testing conducted by

Configuration UII-1H
Configuration I and VI

Blade Radius - In.

Figure 30. Vibratory Moments, Configuration I, VI and UH-1H .

Configuration UH-lil
Configuration II

Blade Radius - In.

Figure 31. Vibratory Moments, Configuration II and UH-

nfiguration UH-1H
nfiguration III

Blade Radius - In.

Figure 32. Vibratory Moments, Configuration III and UH-1H.

g_OI x qi-ui - jusuiojaj XjojBjqiA

50 100 150 200 250 288

Blade Radius - In.

Figure 33. Vibratory Moments, Configuration IV,V and UH-1H.

90 Kt GW = 8500 Lb

Figure 34. Steady Moments - Typical for All Configurations.

318 RPM

qq - aojoj xB3njTjjU93

100 150 200 250 288

Blade Radius - In.

Figure 35. Centrifugal Force vs Blade Radius.

Configuration UH-1H

Blade Radius - In.

Figure 37. Flatwise Stiffness Distribution

rri

Configuration UH-1H
Configuration I and VI
Configuration II

o o o o o o o

• • • • • • <

'O in ^ CO CN

* U I - 9§pg Sujpeog tuojg sduejsiq

Blade Radius - In.

Figure 41. Blade Flexural Axis Comparisons.

Sikorsky Aircraft. In all cases the working endurance limit stresses
shown are reduced from the mean endurance limit by the appropriate
probability and size effect factors. All values shown are at a "no steady"
stress loading condition (R = -1. 0).

The stresses in each of the major blade components were calculated at
four radial positions on the blade. The four stations selected are radial
station 200, 160, 120, and 80 inches. Station 200 is the inboard end of
the constant section of the blade; stations 160 and 120 were selected as
representative midspan stations; and station 80 is at the outboard end of
the root end laminate buildup. At each of these stations, stresses were
determined at several locations on the chordwise cross section. These
locations include the combined flatwise and edgewise stresses at 10 per¬
cent chord, a point of maximum combined stress forward of the 1/4
chord and at the back corner position of the spar where the flatwise
stress is generally a maximum. Maximum edgewise bending moment
induced stresses occur at the trailing edge in the skin and spline mem¬
bers. At each of these locations, stresses were determined for each of
the materials present.

Blade component stresses are determined for the loads shown in Figures
30 to 35 by applying the equation for direct stress, f = P/A, and bending
stress, f = Mc/I (Reference 10). Because the blade structure in each de¬
sign is a combination of materials which have different moduli of elas¬
ticity, the moduli must be considered in the stress calculation. The
equations now become f = PE C /EA and F = McE c /EI. The effect of the
variation in modulus between components is to distribute the load in the
structure in proportion to the modulus of each component. Given the
same position in the structure, a component made of a high modulus ma¬
terial will carry a higher load and have a higher stress level than the
same component made from a lower modulus material. The equations
for the calculation of blade stresses are

f fg + f v

f _ ^cf^c + MfsCyyEic + M e s Cj^ E c

f - Mf C, E + M C F
v r v y y c ev ^xx

F A = F E ^ " f s

(15)

(16)

(17)

(18)

where

f = Total stress at a point on the blade cross
section, lb /in.

f g = Combined steady stress, lb/in.

f = Combined vibratory stress, lb/in.

Pcf = Centrifugal force, lb
Mf s = Steady flatwise moment,in. -lb
M es = Steady edgewise moment, in. -lb
Mf v = Vibratory flatwise moment, in. -lb
M ev = Vibratory edgewise moment, in. - lb

EI XX = Total flatwise bending stiffness, lb/in.

Elyy = Total edgewise bending stiffness, lb/in.

EA = Total axial stiffness, lb

E c = Modulus of elasticity of the component
(material) ,1b/in. ^

Cx X = Distance between the point under consideration
and the chordwise blade neutral axis, in.

Cyy = Distance between the point under consideration
and the neutral axis perpendicular to the chord-
wise axis, in.

= Allowable alternating stress, lb/in.

F f = Endurance limit of material at zero steady
stress, lb /in.

F tu = Ultimate tensile strength of material, lb /in.

For flatwise bending, tension on the bottom side of the beam is consider¬
ed positive. For edgewise bending, tension on the leading edge of the

beam is considered positive.

The stresses obtained for the 90-knot cruise condition for each of the de¬
signs under consideration and the present UH-1H blade are shown in
Tables IV to VIII. The allowable alternating stresses shown in the tables
are obtained from Goodman diagrams for the appropriate materials. A
typical Goodman diagram is shown for 2024T-3 aluminum (Figure 43) to
illustrate the relationship between the steady stress level and the allow¬
able vibratory stress. The margins of safety shown in the taoles are
obtained from the relationship

M. S. =F A /f y -l (19)

where

M. S. = Margin of Safety
F atigue Life Determination

Calculated fatigue life of a rotor blade is a function of the relationship of
the stress level in the component to the allowable stress level (endur¬
ance limit) of the material of which the component is constructed. If the
stress levels in the component were always below the endurance limit
stress, then the component would have an infinite life (assuming no
externally caused damage such as corrosion, foreign objects, etc.).
Design of a component with such low stress levels would be overly con¬
servative however, and would result in a very heavy rotor blade. A
typical rotor blade is designed so that the conditions at which the aircraft
will spend the major percentage of its operating time will not produce
any damaging blade stresses. The effect of those conditions of aircraft
operation which produce damaging stresses is evaluated using the Cumu¬
lative Damage Theory of Miner (Reference 11). Miner’s theory states
that a fatigue crack will be initiated when the summation of the incre¬
ments of fatigue damage equals unity, or

ni/Ni + n 2 /N 2 + n 3 /N 3 +-i^/Nn = 1 (20)

where n n = Number of cycles at a specific stress level

N n = Number of cycles required to initiate a fatigue
crack at that stress level

To calculate the fatigue life of a blade using the cumulative damage theory
we first determine the operating spectrum for the aircraft. A spectrum
consists of each condition of aircraft operation for the typical mission
along with the percentage of total aircraft operating time spent at that

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f'-

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+1 1/5
* r-

— 1 SO

CM CO

co .

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T}i ^-4

■—i id 4

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r- .

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f—1

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m so

00 SO

CM CO

so .uo

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cm m oo

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fc 2

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< ^

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.5 ^ ^ V3

w h fc ^

BLADE STRESS IN LEVEL FLIGHT CRUISE - CONFIGURATION I AND VI
V- 90 Kt GW = 8, 500 Lb

■<*

^-4

CSJ

sO sO

N H

O 00

.uo

. uo

<-l CO

UO .

uo .

* •

PS •

•*

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Tt T*

■*

■s*

-H

oo in

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uo - in
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_u CN
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— ON'
so -
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rP £>P

+1 ^

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O .CO
co in .

-H ^

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O . ■<*
oo m .

m in
* oo as

O CN CO

r- * ■

N N

_u U0>

-H 00 Os

cn cn tJi

CO -

u/2

Stress at Trailing Edge 6,169 * 1,451 5,901 * 1,625 7,996* 1,640 9,027 * 1,532

F Allow 11,950 12,000 11,800 11,750

MS 7.24 6.38 6.20 6.67

TABLE VI. BLADE STRESS IN LEVEL FLIGHT CRUISE - CONFIGURATION II

V=90Kt GW = 8, 500 Lb

i—l

-H

»—4

o tv

ON o

-H

tv o

- CN

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Stress at Trailing Edge 9,281 ± 1,139 8 » 531 ± 1,307 6,600±1,297 6,884 ±1,194

F Allow 11,750 ’ 11,750 11,900 11,850

M S _9 1 32_7^99_8. 18 8. 92

TABLE VII. BLADE STRESS IN LEVEL FLIGHT CRUISE - CONFIGURATION III
_ V = 90 Kt GW = 8,500 Lb

•Nt

T“^

•H

-H

r—^

-H

in .—1

r- cn

in co

_T }1

•N

i-H

•N •

r—< •

r. .

NO CN

<-l CO

'jf

r-T

t-H

•N

i-H

-H

•H

in co

NO CO

O ON

-NO

i“H 00

»• •

t".

H •

•n ■

r—^

nO CM

r -1 CO

•<f co

-'t

<«*

r-N

3 j

O >

CO u, 5

u « s* £

2 0) U f-H c/3

colco o b

t^o

O -CN

'■* f* .

CN »-< ^

^ O

-H *0
^ NO CO

in -cn
o *~* •

no —I no

6 g

■8 »

ctf G

Ss I

*i m [l 3
5 Vh <
C/3 t/3 O b

4—*

co (u
co bo
43 tj
O w

« C Z

^j=i o

-H I'H f—i

.5 Irt rj co
mV* < -

_ ‘-f

co V 3

O T3

« w

_ <50
G £
gS o

G .i-i i—i

s a ^

D,*- 1 ^

C/3 - 1 b

MS 9. 10 7.77 7.69 7.85

VIII. BLADE STRESS IN LEVEL FLIGHT CRUISE - CONFIGURATION IV AND V
_V = 90 Kt GW = 8, 500 Lb

,-h

r-'

i-H

-H

•n

„

f-H

t''

^■1

„

-H

f-H

P“H

•

»-H

*•

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„

»—H

-li

-H

ID CO

O ID

O O'

O' CN

- r-

*—< ,

*« «

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f-H

o .

i-H m

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ID nO

CN r-l

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-L) ID

^ o
o -

CN —i .

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CO O' o

O P

OS »"H

*N •

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* 8
'O ^
N' - ID
ID OO .

u- ;>

j-i

"P u

o “
id ti

Tf W

+J, w

|C0
C !L>

p cd
<

C bfi

o c
•O S
v-( 7-

id ti

tji Co

Hj, m
|to

C 0)

Slii

<*-

r Allow 18,250 19,400 20,700 20,800

MS 2.81 2.53 2.62 2.68

2024 -T3 Aluminum

IS)I - ssarjg XjojBjqiy\

Steady Stress - KSl

Figure 43. Typical Goodman Diagram.

condition. Blade stresses are determined for each of the operating con¬
ditions in the spectrum at a specific location on the blade being analyzed.
Each level of blade component stress is then compared with the S-N dia¬
gram for the appropriate material.

A point anywhere on the S-N diagram represents the existence of some
induced stress for some number of cycles. If the point is on the curve,
it represents a failure, or more exactly the probability that a fatigue

! failure will occur in the number of cycles corresponding to the point.

Several points plotted on a sample S-N diagram (Figure 44 ) illus¬
trate the relation of stress to fatigue life. Point (c ) is below the endur¬
ance limit and therefore contributes noxumulative damage to the com¬
ponent. Point (a) has an abscissa of lCr cycles, while the point on the
curve at that ordinate is at 5 x 10^ cycles. Thus,point (a) represents the
using up of 10^/5 x lCr = .02 of the fatigue life. Similarly, point (b)
represents the using up of 2 x lCr/5 x 10^ = .04 of the fatigue life. If we
assume that these three points represent the stress history of 50 hours of
operation, then the three points represent .06 of the life used up in 50
hours of typical operation. The fatigue life based on the cumulative
damage theory would then be 50 hours/.06 = 833 hours.

Evaluation of the fatigue life of the blade designs under consideration in
this study by the previously described method was not attempted. The
determination of stresses and the frequency of occurrence in maneuvers,
rotor starts and stops and flares are obtainable only from previous test¬
ing of similar model aircraft. For this study, it was not deemed neces¬
sary to perform a rigorous analysis since a relative comparison with the
UH-1H would suffice. Therefore, we chose to calculate only the level
flight cruise condition stresses for each proposed design and for the
present UH-1H design. By comparing each of these stresses with the
material endurance limit stresses and determining their fatigue margins
of safety, we obtain a more valid means of comparing designs with each
other and the UH-1H blade.

Fatigue margins of safety determined for each of the design configura¬
tions are tabulated in Tables IV to VIII. A comparison of fatigue mar¬
gins shows that all designs are at least as good as the present blade.

The metal blade designs including the UH-1H are all very similar in
margin; however, the composite blades show significant increases in
the fatigue margins of safety. Since the mass and stiffness characteris¬
tics of all designs were made similar, the bending deflections would also
be similar in magnitude. Therefore, the strain of the material in each
blade will also be similar. Then the ability of each blade to resist fatigue
damage is only a function of the strain allowable of the material. Strain

allowable is defined as the stress at the endurance limit divided by the
modulus of elasticity of the material. A bar graph of representative

IS^ - SS3J3S ^JOiFjqjA

strain allowables for several typical rotor blade construction materials
is shown in Figure 45 (References 12, 13, and 14). The bar graph illu¬
strates very clearly the advantage to be gained in fatigue life using com¬
posites .

Because of the close similarity between the fatigue margins of the metal
blades and the UH-1H we have assigned the same 2500-hour life as the
UH-1H blade to the metal designs. We have assigned an estimated life
of 5000 hours to the composite designs, an assumption which we feel is
conservative based on the much higher fatigue margins of safety.

SURVIVABILITY ANALYSIS

Since design Configurations I, II and III each had a similar fiberglass
trailing edge structure and each of these had a tubular type spar, only
Configuration I was subjected to a survivability analysis because it was
considered representative of the three designs. Configuration IV con¬
sisted of a carbon trailing edge structure and a twin beam type spar and
was therefore considered sufficiently different from the other designs to
require a separate analysis.

Configurations I and IV were analyzed to determine the amount of damage
or component failure the blade structure could withstand before ultimate
failure. Various types of damage were considered including the complete
separation of the trailing edge spline, cracking of the skin aft of the spar,
a combined spline and skin failure and various bullet holes in the spar
and other structural members. In each case, the structure was analyzed
for loads developed in the 90-knot cruise condition. Two radial stations
were considered, one at 80 inches and the other further outboard at 160
inches. The 80-inch location is just outboard of the root end doubler
buildup area; at this location the edgewise moments are maximum and
the blade cross section is minimum. At 160 inches, the flatwise moments
are maximum and the stress levels for the undamaged blade are also
maximum.

The blade bending moments calculated for the undamaged blade were also
used for the damaged blade analysis. This is because the reduction in
blade stiffness associated with the various damage modes will not change
the undamaged blade bending moments significantly,since the changes in
stiffness occur over only a very small segment of the blade length and
therefore will not result in any significant change in curvature of the
blade. The internal moments on the structure resulting from the local
shifting of the structural centroid (neutral axis)at the damage location are
considered in the analysis. For example, removal of structure from the
trailing edge of the blade shifts the neutral axis forward locally. Centrif-

Vibratory Strain Allowables at R = -1.0

Strain Allowables - * yj n

Figure 45. Material Strain Allowables .

ugal force acting through the mass centroid of the blade produces the
local internal moment at the point of damage. The moment is the cen¬
trifugal force times the offset distance between the mass centroid and
the local neutral axis. Dependent on the type of damage and the particu¬
lar blade configuration, the internal moments can be very significant.

Stresses were calculated at various positions on the remaining undamaged
blade components. In general, the trailing edge of the skin is the most
highly stressed area of that component. Stresses in the spar are a max¬
imum at the aftmost point on the spar contour. In Tables IX and X,
the component stress levels are tabulated for both fatigue stresses and
the peak static stress levels. The margins of safety in fatigue are
shown as a function of the working endurance limit. Margins of safety
for the static stresses are based on the ultimate tensile strength of the
materials.

Damage to the trailing edge spline resulting in complete severing of the
spline member was considered first. Such damage could result from
projectile damage or by a failure of the spline structure. The trailing
edge spline provides the major portion of the edgewise stiffness. The
percentage of total stiffness contributed by the spline increases from the
tip to the root of the blade. For Configuration I, the spline contributes
26% at the tip and 63% at the 80-inch radial station. The spline contrib¬
utes 61% and 80% at the comparable stations for Configuration IV. Since
the spline contributes so much to the edgewise stiffness, it also has a
large influence on the neutral axis of the structure. Loss of the spline
results in a shift in the neutral axis toward the leading edge of the blade
and an internal moment due to centrifugal force which tends to place the
trailing edge in tension.

Examination of the tables shows that complete severing of the spline will
not result in ultimate failure of the rest of the blade structure for either
design concept. The life of the spar in Configuration I would, however,
be limited, as would the carbon skin in Configuration IV. Both compon¬
ents would be life limited on the inboard portion of the blade only.

Failure of only the skin aft of the spar reduces the edgewise stiffness and
to some degree also the flatwise stiffness. However, the contribution of
the skin to the overall blade edgewise stiffness is small compared to the
spline. Since the stiffness contribution of the skin is constant over the
blade length, the percentage reduction in edgewise stiffness lost will de¬
crease going from the tip to the root. Therefore, in the root end region
where the loads are a maximum, the reduction in blade structural prop¬
erties will be a minimum. Tables IX and X again show the results of the
analysis for this failure mode. Both design concepts under consideration
have sufficient life to continue flight based on the positive fatigue margins

TABLE X. STRUCTURAL ANALYSIS FOR VARIOUS MODES OF FAILURE/DAMAGE -

CONFIGURATION IV

CM 00 CM in in CM
CO 00 CM CM OO

— cm cm t'- ^

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m oo h- cm

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cn cn W cn cn —I

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of safety, and both have high ultimate strength margins of safety as well.

A combined failure of the skin and the trailing edge spline was also con¬
sidered where all of the structure aft of the spar is eliminated. The spar
and the remaining skin provide all of the edgewise stiffness. Flatwise
stiffness is not reduced very much since the spline and skin provide very
little of the total flatwise stiffness. The neutral axis will be displaced
toward the leading edge, further forward than when only the spline was
severed. 'The resultant large edgewise mome.t thus produced is reacted
by large tensile stresses in the trailing edge jf the spar. Table IX shows
that the spar in Configuration I is life limite j over its entire spanwise
length. Stresses in the aluminum spar will also exceed the ultimate
strength of the spar inboard of approximately midspan of the blade. The
very high fatigue and ultimate strengths of the composite spar are evident
in Table X. Whereas the aluminum spar stresses would exceed the ulti¬
mate , the composite spar and skin doubler will still have a positive mar¬
gin in fatigue as well as a high ultimate margin of safety.

Ballistic damage to the blade spars in all the design concepts is consid¬
ered survivable. Service experience with aluminum spar blades has
shown that up to . 50-caliber bullet holes in various locations on the spar
are survivable. Figures 46 and 47 are photos of actual in-flight ballistic
damage sustained by Sikorsky main blades. The blade in Figure 46 was
actually flown on a subsequent mission after the . 50-caliber hole was "re¬
paired" with aluminum foil tape. Although no service experience is
available on composite blade spars, we expect that the composite blades
will be even more damage tolerant than the present metal ones. The very
large amount of reserve strength present in the twin beam composite spar,
as shown in the analysis of both the damaged and undamaged rotor blades,
makes this very probable.

X&yttomm

rj^p. * »

I'W”* Twa iiw ««r»

RELIABILITY AND MAINTAINABILITY ANALYSIS

Reliability/maintainability participation in the UH-1H expendable blade
design effort consisted of assisting in the development of R/M related
cost-effectiveness equations and design analysis of baseline and candidate
blade designs. A mathematical model was designed to measure UH-1
cost effectiveness when equipped with the current UH-1 blade and
comparative cost effectiveness values when equipped with each design
candidate expendable blade. Standard nonvariable input values were
supplied by the Government for use in the model. Variable input values
were developed through reliability, repairability, and maintainability
analyses of the baseline UH-1 blade and each candidate blade design.
Reliability/maintainability input variables for use in the model are as
follows:

Input Variable

Units

1. Aircraft down hours

2. Aircraft aborting failure rate

3. Blade mean time between

inherent failures

4. % damaged blades repaired on

aircraft (ORG level)

5. % removed blades repaired off

aircraft (ORG level)

6. % removed blades scrapped (ORG)

7. % removed blades sent to direct

support level

8. % received blades repaired at

direct support level

9. % received blades scrapped at

direct support level

10. % received blades repaired at

depot level

11. MMH/inspection - on aircraft (ORG)

12. MMH/repair - on aircraft (ORG)

13. MMH/repair - off aircraft (ORG)

14. MMH/repair - direct support

15. GSE cost - on aircraft repair (ORG)

16. GSE cost - off aircraft repair (ORG)

17. GSE support cost (ORG)

18. GSE cost per repair - direct support

19. GSE support cost - direct support $/aircraft

20. GSE cost per repair - depot $/repair

Down hours/flight hour
Aborting failures/flight hour
Blade hours

MMH/blade
MMH/repair
MMH/repair
MMH/repair
$/repair
$/repair
$/a ire raft
$/repair

Input Variable

Units

21.

GSE support cost - depot

$/ aircraft

22.

Parts /material cost - on aircraft
repair - ORG level

$/ repair

23.

Parts/material cost - off aircraft
repair-ORG level

$ / repair

24.

Parts /material cost - direct support

$/ repair

25.

Blade overhaul cost - depot

$/blade

MATH MODEL INPUT VARIABLES - DISCUSSION AND DEFINITION

Aircraft Down Hours

The influence of the baseline UH-1 main rotor blade and of each candi¬
date blade design upon aircraft downtime was estimated and treated as
an input variable to the math model for use in measuring changes in
aircraft operational availability relative to each blade design. Baseline
UH-1 down hours per flight hour were calculated in the following
manner:

Operational Availability =

Flying time + Ready time
Total time available

( 21 )

where operational availability = 75% (value supplied by USAAMRDL)
utilization rate = 41 hours/month/aircraft

time available = 720 hours/month

Flying time + ready time = (720) (. 75) = 540 hours/month
Total downtime = Time available - (Flying time 4-

ready time)

= 720 - 540 (22)

= 180 hours/month

town hours/Flight hour = . 4. 38 < 23 >

The 4. 38 down hour per flight hour arrived at above is considered to
include supply, administrative, and maintenance downtime and
represents the UH-1 DH/FH value used as a baseline value for purposes
of this study.

Aircraft Aborting Failure Rate

The aircraft aborting failure rate was originally introduced to the math
model to measure variations in mission reliability with changes in main
rotor blade designs and resultant impact on cost effectiveness. In

order to effectively use this parameter in the math model a baseline
value for the UH-1 aircraft and the UH-1 main rotor blade aborting
failure rates was required. The actual values for the UH-1 aircraft
proved to be unavailable,and therefore a representative value was esti¬
mated for the overall aircraft using Sikorsky in -house aborting failure
rate data adjusted to reflect the UH-1 configuration. The estimated
value assigned to the UH-1 was . 015, or 15 mission aborting failures
per 1000 flight hours. Due to the lack of background data relating speci¬
fically to UH-1 main rotor blade aborting failure rates and their
influence on the overall aircraft aborting failure rate, mission reli¬
ability was treated as a constant for all blade designs throughout the
study.

Blade Mean Time Between Inherent Failures

The MTBR value assigned to the UH-1 baseline blade for inherent causes
was taken directly from Reference 2 arid is cited as 3,733 blade hours.
The rationale connected with the use of this value is discussed in the
following paragraphs.

Blade Retirement Life

Baseline UH-1 blade retirement life was cited as 2500 blade hours in
Table XXXXIV. Retirement life values cited for candidate expendable
blade designs were based upon load and stress and analyses conducted
for each of the proposed designs.

Remove-Repair-Scrap Percentage Values

Math model input values for these parameters were calculated directly
from the candidate blade repairability analyses which are presented in
Appendix II. Baseline values for the UH-1 blade were furnished by
Table XXXXIV.

Maintenance Man-Hour Values

All man-hour input variables to the math model were taken from the
maintenance task analyses conducted for each blade repair procedure.
Refer to the maintainability analyses presented on page 121 for the
methodology used to calculate maintenance man-hour values.

Support Equipment Cost Factors

Support equipment cost factors were computed on the basis of (1) the
cost of special support equipment required to support 24 aircraft per
site and (2) the cost of special support equipment per repair action.

The cost values cited represent only that cost incurred for special
support equipment over and above that which is already in existence for
the current UH-1 blade. A typical calculation for the cost of support
equipment at the direct support level of maintenance follows:

1. Cost of SSE per 24 aircraft = $15,304. 00

2. Frequency of repair at DS level

requiring the use of SSE = 3,820 Blade Hours

3. 24 aircraft per site per 10-year

life cycle = 240,000 Blade Hours

Cost per A/C =

Cost of SSE per site
No. of A/C per site

$15,304
“24-

= $637. 00 (24)

Cost per repair^ =

Cost of SSE per Site

(Blade hrs per life cycle 24A/C)HFreq repair atDS)

$15,304 00
240(10)^/3,820

= $243. 00 per repair

(25)

Parts/Material Cost

Parts/material cost values were computed on the basis of dollar cost of
parts and material per average repair procedure at each level of main¬
tenance. The cost of repair kits containing all required materials for
minor and/or extensive repair of fiberglass or carbon skin damage was
calculated. Each kit contains required parts as well as materials for
accomplishing all fiberglass and carbon repairs. Refer to Appendix 11
for fiberglass and carbon repair procedures.

Blade Overhaul Costs

Depot level overhaul costs were calculated for each candidate blade
design based upon part, material and labor cost estimates. Support
equipment cost relative to depot overhaul is not included in the cited
values since Table XXXXIV referred this parameter as zero for both the
UII-1 and candidate blade designs.

100

order to effectively use this parameter in the math model a baseline
value for the UH-1 aircraft and the UH-1 main rotor blade aborting
failure rates was required. The actual values for the UH-1 aircraft
proved to be unavailable,and therefore a representative value was esti¬
mated for the overall aircraft using Sikorsky in-house aborting failure
rate data adjusted to reflect the UH-1 configuration. The estimated
value assigned to the UH-1 was . 015, or 15 mission aborting failures
per 1000 flight hours. Due to the lack of background data relating speci¬
fically to UH-i main rotor blade aborting failure rates and their
influence on the overall aircraft aborting failure rate, mission reli¬
ability was treated as a constant for all blade designs throughout the
study.

Blade Mean Time Between Inherent Failures

The MTBR value assigned to the UH-1 baseline blade for inherent causes
was taken directly from Reference 2 and is cited as 3,733 blade hours.
The rationale connected with the use of this value is discussed in the
following paragraphs.

Blade Retirement Life

Remove-Repair-Scrap Percentage Values

Maintenance Man-Hour Values

DEVELOPMENT OF RELIABILITY INDEX

Data Source

The presented data was reviewed and the reported failure modes and fre¬
quencies were apportioned to the UH-1 component parts. The mean time
to removal values quoted in Table D1 of Reference 2 were converted
to mean time between removal values for use in determining total
inherent and external failure rates for candidate expendable blade
designs. Background data relating to the reliability and maintainability
of fiberglass skins was taken from contractor experience with the
Sikorsky improved rotor blade currently being tested on the CH-53
helicopter, Sikorsky experience with fiberglass fuselage skin panels,
service experience with fiberglass propellers, and results of Sikorsky
in-house testing relative to the abrasion resistance qualities of fiber¬
glass.

Reliability Values

Table XI is a compilation of published UH-1 main rotor blade mean times
to removal due to various failure modes and their conversion to mean
times between removal. Mean time to removal is defined as the sum
of the times at removal for all blades divided by the number of blades
removed, or

i = n

where ti = the total time at removal of the ith blade in hours
n = the number of blades removed

Reliability data presented in MTR form is not suitable for predicting
blade life-cycle failure occurrences or blade life-cycle inventory
requirements. Values computed as shown above do not reflect total
time generated by the entire blade population. They represent only
those hours recorded on removed blades at the time of removal and
therefore result in values which are considerably lower than those
which should be used for logistics purposes or for determining blade
life-cycle cost. In order to account for total blade population and to
utilize the published data for prediction purposes the MTR's of Table
D-l of Reference2 were converted to MTBR's and are reported in
1 able XI.

101

Support Equipment Cost Factors

1. Cost of SSE per 24 aircraft = $15,304. 00

2. Frequency of repair at DS level

requiring the use of SSE = 3,820 Blade Hours

3. 24 aircraft per site per 10-year

life cycle = 240,000 Blade Hours

Cost per A/C =

Cost of SSE per site
No. of A/C per site

$15,304

~n —

= $637. 00 (24)

Cost per repair™ =

Cost of SSE per Site

(Blade hrs per life cycle 24A/'C)r(Freq repair atDS)

_ $15, _ 3 04. 00 _ £243.00 per repair (25)

240(10) /3,820

Parts/Material Cost

Blade Overhaul Costs

100

DEVELOPMENT OF RELIABILITY INDEX

Data Source

Reliability Values

i = n

where tj = the total time at removal of the ith blade in hours
n = the number of blades removed

Reliability data presented in MTR form is not suitable for predicting
blade life-cycle failure occurrences or blade life-cycle inventory
requirements. Values computed as shown above do not reflect total
time generated by the entire blade population. They represent only
those hours recorded on removed blades at the time of removal and
therefore result in values which are considerably lower than those
which should be used for logistics purposes or for determining blade
life-cycle cost. In order to account for total blade population and to
utilize the published data for prediction purposes the MTR’s of Table
D-l of Reference 2 were converted to MTBR's and are reported in
Table XI.

101

Inherent Damage

The established values of Table XI were used to construct a reliability
profile of the baseline UH-1 blade. Blade inherent failure modes were
determined and quantified in terms of MTBR based upon the number of
failure occurrences given in Table XI divided into 6, 698,706 total blade
hours. This total blade hour value was arrived at by using a represen¬
tative mean time between removal value of 914 hours, as selected from
Table E-lof Reference 2, multiplied by the total number of occurrences
of Table XI. The MTBR values established through the above procedure
were then allocated to the component parts of the UH-1 blade. For
example,

Table XI indicates that 400 removals took place due to "BONDING
SEPARATION".

MTBR = — 69 | 8 qq 06 = 1 P er 16,747 blade hours = . 000060

bond separation (27)

The MTBR value was then allocated to UH-1D blade component parts
which are susceptible to "BONDING SEPARATION".

Component Parts

Apportionment

Abrasion Strip Bonding Separation =

Core to Spar Bonding Separation =

Skin to Core Bonding Separation
Doubler Bonding Separation =

Trailing Edge Strip to Core S =

Trim Tab in Bonding =

Skin to Spar Separation =

Skin to Trailing Edge Strip Separation =

. 000018
. 000004
. 000019
. 000004
. 000002
. 000004
. 000004
. 000005

. 000060

Apportionment values were established by reference to various data
sources including Sikorsky in-house reliability data relative to main
rotor blades. The MTBR values thus established and apportioned for tie
entire UH-1 blade served as the basis for the evaluation of the inherent
failure modes of all candidate blade designs. The established MTBR
values were adjusted to compensate for basic differences in candidate
blade designs. For example, the UH-1 blade aluminum skin MTBR was
established at 43 removals per 10^ blade hours. This same value was
assigned to the skin of our candidate aluminum extruded spar design
because there is no basic difference between the two skin designs. How¬
ever, the same extruded aluminum spar design utilizing a fiberglass

102

103

104

TABLE XI ■ Continued

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Heat Damage 17 349.7 394,042

Blistered 10 394.4 669,870

Burned 6 238.4 1,116,451

Heat damage 1 504.0 6,698,706

106

skin rather than aluminum exhibits a MTBR for skin problems of 126
removals per 10” blade hours. This adjusted value reflects additional
failure modes inherent in fiberglass materials as compared to aluminum
(see Table XXXII, Appendix II).

External Damage

The frequency of removals caused by external damage is treated as a
constant throughout the study as directed by USAAMRDL. The given
mean time to removal of 400 hours was converted to a mean time between
removal of 1,211 blade hours (excluding no-failure causes and unknown
causes). The distribution of UH-1 blade externally caused damage modes
is presented in Table XI.

While the external damage rate is a constant value for all blades con¬
sidered in the study, it is necessary to determine the changes in blade
external damage repairability resulting from differences in the inherent
design characteristics of each candidate blade. This was accomplished
by allocating the MTBR value for external damage to the component parts
of each candidate blade design on a percentage basis. For example:

Battle damage - 120 occurrences per l(fo blade hours for the UH-1 blade.
Candidate blade planform is the same as for the UH-1 blade, and there¬
fore it is reasonable to assume that the external damage pattern for the
candidate blades will be the same as for the UH-1 blade, and all candi¬
date blades could also be expected to experience 120 incidents of battle
damage per 10^ blade hours. In order to determine how many of these
incidents will occur in a given blade component part, the percent of total
blade surface area occupied by the part is calculated and multiplied by
the battle damage incidence rate of 120 per 10^ blade hours. This pro¬
cess is repeated for all external damage modes and all component parts
until the entire external damage rate (826 occurrences per 10” blade
hours) is allocated.

The reliability apportionment of the baseline UH-1 blade inherent damage
rates is presented in Table XII. External damage rates for the baseline
blade are not apportioned since it was not necessary to calculate the
degree of repairability of the UH-1 blade. Baseline UH-1 blade repair-
ability was determined from the published data provided by USAAMRDL.
External damage rate apportionment for all candidate blade designs is
presented in the repairability analysis of Appendix II.

107

TABLE XII. RELIABILITY APPORTIONMENT-BASELINE UH-1D BLADE

I, Inherent Damage

Frequency of
Occurrence per

Blade Component

Failure Mode

10 ^B ladeHours

1. Spar

A. Bonding separates

from core

4.0

B. Elongation of bush-

ing holes

10.0

C. Cracks

D. Abrasion strip sep-

14.0

aration

18.0

E. Corrosion

F. ILtted, abraded or

9.0

eroded abrasion
strip

12.0

67.0

2. Core

A. Bonding voids

20.0

B. Water contamination

10.0

30.0

3. Skin (Aluminum)

A. Unbonded at leading

or trailing edge

9.0

B. Corrosion

2.0

C. Cracks

32.0

43.0

4. Retention Bushings

A. Cracks

10.0

B. Wear

9.0

C. Corrosion

2.0

21.0

5. Doublers (includes

A. Bonding separation

4.0

grip and drag plates)

B. Corrosion

2.0

C. Cracks

10.0

16.0

108

TABLE XII - Continued

I. Inherent Damage - Continued

Blade Component

Failure Mode

Frequency of
Occurrence per

10° Blade Hours

6. Trailing Edge Strip
(aluminum)

A. Bonding Separation

B. Cracks

2.0

10.0

12.0

7. Trim Tab

A. Loose Rivets

1.0

B. Unbonded

4.0

5.0

8. Counterweights

A. Loose

B. Corroded

1.0

2.0

9. General

72.0

Total Inherent Damage

268.0

II. Total External Damage

826.0

III. Total Blade Damage

1,094.0

109

FAILURE MODE AND EFFECTS ANALYSIS

A failure mode and effects analysis was conducted for each of the candi¬
date blade designs and is presented in Appendix II. Possible failure
modes anticipated in the blades’operational environment, their effect
upon the blades’functional capability, and probable symptoms and
methods of detection were investigated. Specific design features incor¬
porated to minimize and/or reduce the effect of anticipated failure modes
are as follows:

Configuration I and VI

1. Use of 6061 aluminum in spar to decrease crack propagation
and reduce corrosion.

2. Heavy wall spar to resist external damage and increase field
repairability by permitting more extensive blend out procedures.

3. Use of fiberglass skin to increase skin fatigue life and blade field
repairability.

4. Attaching point bushings replaceable in the field.

5. Increased bonding area and use of 6061 aluminum in grip pad, drag
plate, and doublers provide greater margin of safety, increased
crack propagation time and increased corrosion resistance.

6. Use of nonmetallic honeycomb core provides increased elastic

memory and reduces effect of external damage on honeycomb
(Configuration I only).

7. Trailing edge spline corrosion free and repairable.

Configuration II

1. Three-piece spar construction offers increased redundancy.

2. Stainless steel leading edge provides greater erosion resistance
and increased durability.

3. Use of fiberglass skin to increase skin fatigue life and blade field
repairability.

4. Attaching point bushings replaceable in the field.

5. Increased bonding area and use of 6061 aluminum in grip pad, drag

110

plate, and doublers provide greater margin of safety, increased
crack propagation time, and increased corrosion resistance.

6. Use of nonmetallic honeycomb core increases elastic memory and
reduces effect of external damage on honeycomb core.

7. Trailing edge spline corrosion free and repairable.

Configuration IV and V

1. Twin beam concept provides a potentially fail-safe spar design with
redundant load paths. Use of unidirectional fiberglass in beam
construction provides greater fatigue life, slow crack propagation,
and greater repairability characteristics than other concepts.

2. Leading edge protective coating is field replaceable.

3. Carbon skin increases skin fatigue life and repairability in the
field.

4. Attaching point bushings are field replaceable.

5. Increased bonding area and use of 6061 aluminum in grip plate,drag
plate, and doublers provide greater safety margin, increased
crack propagation time, and increased corrosion resistance.

6. Use of nonmetallic honeycomb core provides increased elastic
memory and reduces effect of external damage on honeycomb.

7. Trailing edge spline corrosion free and repairable.

RELIABILITY SUMMARY

The MTBR'j' shown in Tsble XIII for die candidate blade designs reflects
removal rates in excess of those attributed to the baseline UH-1 blade.

It must be remembered that the influence of design changes upon the
external damage rate is not reflected in these values. For this reason
the values shown must be considered conservative.

The use of fiberglass skins in all candidate designs is partially respon¬
sible for the higher removal rates shown for these blades,since the in¬
herent failure modes of fiberglass appear to be greater than those for a
comparable aluminum skin. This is offset, however, by the vast in¬
crease in field level repairability which is afforded by the use of fiber¬
glass and the subsequent reduction in the blade scrappage rate. The

111

*w»sr*ir »%

TABLE XIII. FAILURE RATE SUMMARY

Blade Design

Mean Time Between Removal

Mean Time To Removal

Inherent

External

Total

Inherent

External Total

Baseline

UH-1

3,733

1,211

914

547

400

442

Configuration

I and VI

2,680

1,211

833

513

400

436

Configuration

2,850

1,211

850

502

400

450

Configuration

IV and V

1,200

1,211

603

498

400

435

contractor feels that the candidate blade values shown would be sub¬
stantially higher had the impact of design changes on the external damage
rate been a factor in the study. Also.no credit is taken for solving debond¬
ing problems. It is fully expected that the mean time between removals
of these configurations will be higher than the Bell blade. However, for
conservatism, the above hours will be used.

MAINTAINABILITY ANALYSIS

Initial analysis of the cost effectiveness of the baseline UH-1 blade
indicated that the poor repairability characteristics inherent in the
design and the subsequent high scrappage rate were the primary factors
contributing to the high life-cycle cost of the blade. The obvious solu¬
tion to this problem is to produce a blade with a comparable design life
but with a recurring cost so low that repair becomes more costly than
replacement. The second best approach is to produce a blade which
lends itself to cost effective repair by increasing the repair incidence
of failure modes which formerly caused the blade to be scrapped or
returned to depot.

The maintainability portion of this study deals with the second of these
solutions.

Candidate Blade Repairability

The extensive use of fiberglass in all three candidate blade designs
makes possible significant increases in repairability with respect to the
baseline blade. Extensive repair of large blade areas can be accom¬
plished by the use of prefabricated blade sections and prestocked

112

repair kits containing all required repair materials. Bonding agents
which are room temperature curable eliminate the need for expensive
and cumbersome heat treating equipment,thereby providing increased

latitude in assigning repair levels and enhancing unit self-sufficiency in
the field. Blade inspection procedures above and beyond those which are
currently in use should not be required. Existing balancing and tracking
equipment will be compatible with all candidate blades; however, balanc¬
ing technique and procedure may become more significant in view of the
extensive field repairs now possible.

Repairability Analysis

Each candidate blade design was subjected to a repairability evaluation
based upon the failure mode and effects analysis previously conducted.
Repairability analyses are presented in Appendix II . A scrap versus
repair decision was made for each failure mode based upon the follow¬
ing general criteria.

Scrappage Causes

1. Extensive deformation of spar wall or leading edge.

2. Cracked or punctured spar.

3. Extensive bonding separation of spar closure piece or
core to spar bond.

4. Externally caused skin damage which extends into spar
wall, transition doublers or trailing edge spline.

5. Extensive bonding separation or cracks in trailing
edge spline.

6. Extensive delamination or bonding separation of grip
doublers, drag link doublers, or transition doublers.

7. Extensive damage at root end attaching points.

8. Overstress conditions such as overspeed, warpage, or
other blade deformation.

The causes for scrappage cited above are general in nature and vary
depending upon individual candidate blade design. The twin beam spar
concept, for example, can tolerate greater damage and exhibits a great¬
er degree of repairability with respect to spar damage than do the other

113

candidates. Refer to the repairability analyses of Appendix II for candi¬
date blade scrap/repair description.

Level of Repair Decision

Level of repair decisions were made concurrently with the scrap/repair
decision for each candidate blade. Repair levels were assigned on the
basis of feasibility in terms of aircraft downtime, degree of skill required,
and support equipment and facilities required. All fiberglass repair pro¬
cedures are designated as direct support level repair procedures and are
considered to be within the capabilities of the average aircraft rotor and
propeller repairman (MO 68 E 20). Individual training in repair technique
and procedure will be required. Tables XIV and XV summarize candidate
and baseline blade repairability and levels of maintenance.

ITABLE XIV. BLADE REPAIR ABILITY AND LEVEL OF MAINTENANCE

Blade Design

% Repairable
QRG D S Depot

BStiif'. ^

Total

Repair_Scrap

3aseline UH-1

12%

19%

0 30%

39%

31%

69%

Configuration

I and VI

44%

33% H%

54%

46%

Configuration II

44%

33% 12%

53%

47%

Configuration

IV and V

74%

13% 8%

77%

23%

■■■■■■£!

mHsmwmwnmamawm

Blades per Million Hours

Percent of Removed

Blade Design

Removed

Repaired

Scrapped

Blades Repaired

Baseline UH-1

1094

339

755

31%

Configuration

I and VI

1201

646

555

54%

Configuration II

1177

624

553

53%

Configuration

IV and V

1659

1288

371

77%

114

Repair Procedures (Refer to Appendix II for detailed procedures.)

Eight repair procedures which can be considered peculiar to the candi¬
date blade designs are outlined in the following pages. These procedures
will be used to repair blade damage in the field which previously would
have caused the blade to be scrapped or returned to the depot for exten¬
sive repair. Figures 48 through 51 illustrate types of damage which
are not now field repairable. General repair procedures such as
corrosion removal and treatment, repair of loose or missing hardware,
etc. , are considered standard or commonplace and are not included.

Allowable field repair procedures peculiar to candidate blade designs
are as follows:

1. Leading edge polyurethane coating restoration (Configuration IV)

2. Leading edge repair (Configuration IV)

3. Leading edge blend repair (Configuration I)

4. Twin beam repair (Configuration IV)

5. Fiberglass skin repair - patch (Configurations I, II & IV)

6. Fiberglass skin repair - plug (Configurations I, II, & IV)

7. Attaching point bushing replacement (Configurations I, II, & IV)

8. Fiberglass trailing edge spline repair (Configurations I, II, & IV)

Blade Leading Edge Polyurethane Coating Repair - The polyurethane
coating used to protect the leading edge of the blade can be restored in
the field as required. Repair capability covers the full range of
restoration from repair of minor pitting to complete stripping and
reapplication of the coating to the entire leading edge surface. The
materials required to accomplish the full range of restoration can be
supplied in kit form. No special support equipment is required and no
special skill or experience is required.

Leading Edge Carbon Repair - Damage to the leading edge of the
"twin beairT spar configuration such as dents, gauges or penetration is
repairable in the field provided the damage is not extensive enough to
have penetrated to the twin beam and honeycomb core. Prefabricated
sections of leading edge consisting of 6-inch, prescarfed, ready-to-
install nose blocks can be supplied in kit form. The repair procedure
will consist of cutting out the damaged area and replacing it with the
part provided in the kit. A scarfing tool provided by the manufacturer
will be available for reworking the leading edge adjacent to the cutout
section for mating with the prefabricated replacement. A simple
fixture will be required to apply pressure to the replacement insert
during the bonding cure period. This fixture can be supplied by the
manufacturer or fabricated locally. This repair procedure should be
accomplished in a sheltered area and will require a degree of skill
acquired through special training which can be given in the field.

115

Figure 50. Blade Gash Damage.

Figure 51. Blade Dent Damage.

117

Leading Edge Blend Repair - This repair procedure will allow rework of
the extruded aluminum spar leading edge to blend out damage resulting
from external causes up to 1/4 inch in depth. The repair technique is
standard for this type of repair and requires no additional skill beyond
that of a trained metal working technician. No special support equipment
or facilities are required.

Fiberglass Skin Repair - Patch - Fiberglass skin damage such as abra¬
sion, delamination, bonding separation, crazing and blisters can be
repaired by simply removing and replacing the affected area. Repairs of
this type are not limited by size of affected area and have been success¬
fully accomplished and flight tested at Sikorsky Aircraft. This procedure,
with variations, is also applicable to external damage which penetrates
the skin and extends into the honeycomb core. The repair procedure illus¬
trated in Appendix Il(Figure 59 through 72)was actually performed on a
recent flight test article to repair extensive skin to honeycomb bonding
separation which occurred as a result of damage to the trailing edge
caused by contact with a foreign object. All required materials, includ¬
ing skin patches up to 1 square foot in area, can be furnished in kit form.
Skin material required for patches exceeding 1 square foot can be supplied
in bulk form. The repair should be performed in a sheltered area free
from environmental influence and requires the use of support equipment
in the form of a compression blanket. A high degree of skill is not re¬
quired; however, training in the repair technique will be necessary.

Fiberglass Skin Repair - Plug - This procedure is used in conjunction
with the patching procedure to repair damage which extends through the
skin and either penetrates the entire blade or causes extensive damage
to a large volume of honeycomb core. Prefabricated skin/core sections
of standard or varying sizes can be furnished in kit form to replace
damaged sections of the skin and core. Figures 73 through 84 of Appen¬
dix II illustrate this type of repair, with the patching procedure, this
repair should be performed in an area free of environmental influence
and will require the use of a compression blanket. Again, training in the
repair technique will be required.

Attaching Point Bushing Replacement - Removal and replacement of worn
attaching point bushings can be accomplished in the field through the use
of a dual bushing arrangement. Standard techniques for pressing out and
inserting steel bushings can be utilized. Skill levels above and beyond
those which are presently available will not be required.

Twin Beam Repair - Damage to the unidirectional fiberglass twin beam
is also repairable in the field. Foreign object damage which penetrates
the skin and enters the beam can be repaired by removing the damaged

118

skin and beam section and replacing with tapered unidirectional pre-
molded patches. A router and a router template, which would be
furnished by Sikorsky Aircraft, will be required to accomplish the
repair. No exotic special support equipment is required. Training in
the repair technique will be necessary although a great deal of skill is
not required. This repair should be accomplished in a sheltered area
free from environmental influence.

Trailing Edge Spline Repair - Damaged sections of the trailing edge
spline can be rep ired ; n a manner similar to that used for the carbon
leading edges. Prefabricated, prescarfed sections of trailing edge spline
can be supplied. The repair is accomplished by removing the damaged
section and replacing it with the prefabricated replacement. Bonding
and compression techniques similar to those used in the patch and plug
repair procedures will be utilized.

Repair Kit - A single repair kit containing all the required materials
for the fiberglass and carbon repairs detailed in Appendix n can be
assembled. Repair kit contents will be as follows:

a. 1-square-foot fiberglass or carbon skin panels

b. 1-square-foot fiberglass or carbon skin/honeycomb core
prefabricated sections

c. 6-inch prefabricated, prescarfed leading edge nose blocks

d. 6inch prefabricated, prescarfed trailing edge spline sections

e. 180-240 grit sand paper

f. MEK solvent

g. Adhesive stripper

h. Cotton and rubber gloves

i. Masking tape

j. Mixing cup, wooden spatula, serrated spreader

k. Teflon film

l. Nose template

m. Scrim cloth

n. Scarfing tools

Inspection Procedures

Inspection procedures adequate to detect each mode of failure to which
the candidate blades are susceptible will be similar to those which are
currently in use for the UH-1 blade. In addition, tolerance checks can
be performed at predetermined points along the blade span to check for
subsurface delamination and bonding separation of the twin beam. Skin
defects and skin to core bonding separation will be detectable visually
and by tapping and pressure procedures.

119

Balance and Tracking Procedures

All Sikorsky blades are interchangeable individually or as a set. Any
of the six configurations in the report can also be made to be interchange¬
able. Interchangeability would commence by closely controlling the
weight and mass distribution of the blade components during the fabri¬
cation stage. The dimensions and tolerances of the component parts of
the blade would be held to specified limits so that each blade at final as¬
sembly would fall within a specified spanwise moment tolerance. The
spar mass distribution would be controlled during the fabrication stage
by checking its weight and spar moment. The trailing edge spline weight
would be especially controlled because of its extreme location from the
chordwise center of gravity. The trailing edge skins are thin and should
represent no problem because they are light and would have very little
fluctuation in weight. The counterweight package (whether integral or
molded) would be determined during the design phase and checked out
during assembly of the first few blades. Any changes would be minor
and would be done at this time. Static balance is performed on a balance
stand with a master blade. The spanwise moment of each blade is
matched by inserting tip weights just inboard of the tip cap. By closely
controlling weight as described above, the spanwise moments of Table I
(which are approximately 27,000 in.- lb.) could be held to ± 6 in.- lb. per
blade which is extremely close for balance. Aerodynamic and dynamic
balancing is accomplished on the Sikorsky 2000 hp main rotor test stand
by adjusting the blade pitching moment and track characteristics to those
of a master blade. Aerodynamic balancing consists of adjusting external
trailing edge trim tabs to match the blade pitching moments at low angles
of attack. Dynamic balancing entails matching the blade pitching moments
and track at high collective pitch angles by chordwise adjustments to the
blade tip and root end weights. The blade pitch moments are obtained by
measuring the steady loads in the rotor head rotating control rods modi¬
fied by the addition of force load cells. Track measurements are ob¬
tained using a Chicago Aerial Electronic Blade Tracker.

Because the blade pitching moment and track characteristics are matched
on a whirl stand, no further adjustments are required to the blades when
they are installed on an aircraft. Only two installation adjustments are
required - rotor trammelling and tracking. The rotor assembly is moun¬
ted on a fixture for alignment of the blade tips by adjustment at the rotor
head drag struts. Tracking is accomplished after rotor installation. Any
suitable tracking device such as the Chicago Aerial Electronic Blade
Tracker, the Chadwick-Helmut Strobex Tracker, or a flag, may be used.
The only adjustments required are to the rotating push-pull rods to put
the blades in track at normal rotor speed and a moderate pitch angle.

Since the blade pitching moment and track characteristics have been

120

previously matched on the whirl stand, they will stay in track throughout
the rotor speed, power and airspeed range of the helicopter and no fur¬
ther trim tab or chordwise balance weight adjustments are necessary.

Maintenance Man Hour Per Flight Hour

The maintenance man hour per flight hour values cited below represent
the estimated MMH/FH at the organizational and direct support levels of
maintenance. The values include inspection, diagnosis, repair, and
checkout time. Cure time and time to secure replacement parts and ma¬
terials are not included. The values are obtained by determining the
weighted average man-hours to repair at each maintenance level and divi¬
ding by the frequency at which the repair actions occur.

MMH/FH

Configuration I

Organizational Level

A. Inspect/repair on aircraft

B. Inspect /remove and replace/disposition

.0002

.0195

Total ORG

.0197

Direct Support Level

A. Inspect and repair

B. Disposition (scrap or return to depot)

.0040

.0006

Total Direct Support . 0046
Total Configuration I MMH/FH - . 0243

MMH/FH

Configuration II
Organizational Level

A. Inspect/repair on aircraft .0001

B. Inspect/remove and replace/disposition .0194

Total ORG .0195

121

Configuration II - Continued

MMH/FH

Direct Support Level

A. Inspect and repair

B. Disposition (scrap or return to depot)

.0047

.0006

Total Direct Support

.0053

Total Configuration II MMH/FH

.0248

Configuration IV

MMH/FH

Organizational Level

A. Inspect and repair on aircraft

B. Inspect/remove and replace

.0000

.0292

Total ORG

.0292

Direct Support Level

A. Inspect and repair

B. Disposition (scrap or return to depot)

.0189

.0006

Total Direct Support .0195

Total Configuration IV MMH/FH = .0487
Man-Hour Allocation-Allowable Field Repairs

Active Man-Hrs

1. Leading edge polyurethane coating restoration

touch-up 1.5

full restoration 8.0

2. Leading edge repair-composite 8.0

3. Leading edge blend repair-aluminum 1.5

4. Twin beam repair 16.0

5. Fiberglass skin repair-patch 8.0

6. Fiberglass skin repair-plug 16.0

7. Attaching point bushing replacement 6.0

8. Fiberglass trailing edge spline repair 8.0

122

COST-EFFECTIVENESS ANALYSIS

The cost-effectiveness model is used to evaluate the cost effectiveness of
the UH-1 aircraft equipped with the baseline blade and each of the candi¬
date rotor blade designs. As shown by TableXVI, Configurations I and
II both yield more cost effective aircraft than the baseline blade ;
Configuration I has a slight edge. Configurations III and IV are less
cost effective than the baseline blade. The cost effectiveness differences
seem relatively small until translated into an equivalent dollar measure.
This measure, fleet effective cost, is defined as the fleet life-cycle cost
of N' aircraft equipped with a candidate rotor system design where N' is
fleet size adjusted to maintain the fleet effectiveness of 1, OCX) UH-1 air¬
craft equipped with the baseline blade. In this manner, any difference in
aircraft mission effectiveness can be translated into an equivalent fleet
life-cycle-cost increment.

As shown by Table XVII,the cost differences on a fleet basis are signifi¬
cant. For example, the most cost effective design, Configuration I, can
save over $ 12 million for a baseline fleet of 1,000 aircraft. Similarly,
half of this saving is available for a 500 aircraft fleet and 50% more can
be saved for a 1, 500 aircraft fleet.

None of the blade configurations are truly expendable, since none become
more cost effective with the elimination of depot repair. In all cases,
the cost of replenishing the system with additional spare blades is greater
than the cost savings realized by eliminating depot repair costs. How¬
ever, both Configurations I and II can be treated as expendable and still
be significantly more cost effective than the baseline configuration. For
example, Configuration I saves $12.09 million over the baseline con¬
figuration. If depot level repair is eliminated, $11.64 million is still
saved. Some considerations beyond the scope of this cost effectiveness
analysis can possibly make this direction worthwhile. For example, if
near-expendability can be achieved on a number of aircraft components,
it may be possible to reduce the extent of depot facilities required or the
indirect burdens of maintenance management.

The relatively low cost effectiveness of Configurations III and IV is due
primarily to the use of high cost advanced technology materials. Since
materials currently used in blade manufacture tend to increase in cost
with time and advanced technology materials tend to decrease, the rela¬
tive value of these rotor blade designs may shift in future applications.
Each configuration is reanalyzed for the 1980 time period using the
following assumptions:

1. The UH-1 aircraft is assumed to exist as the baseline vehicle in
1980 at no change in cost other than that generated by the blades.

123

L2<

- k *o*. «r v *wr- ,, ?*r

2. Engineering, manufacturing, and maintenance labor rates, in¬
cluding overheads, are assumed to increase by 50% relative to
the 1972 time period.

3. The costs of aluminum and steel are assumed to increase by 30%
due primarily to increased labor costs in the materials supply
industries.

4. In opposition to the cost increase in (3), unit cost savings due to
increased material production and availability are applied to the
acquisition cost of advanced technology materials.

5. A decrease in manufacturing labor time is applied to configurations
with advanced technology materials to account for learning of ad¬
vanced manufacturing techniques.

6. Configuration IV is slightly modified for 1980 and re-identified as
Configuration V.

With these assumptions, the cost effectiveness analysis yields the values
shown in Tables XVIII and XIX for 1980.

For the 1980 time period, Configurations III and V have joined Config¬
urations I and II in providing greater aircraft cost effectiveness than the
baseline blade. Configuration V displaces Configuration I as the most
cost effective blade design. The fleet effective cost saving of Configura¬
tion I over the baseline is almost $17 million, which is more pronounced
than in 1972. The most cost effective blade for 1980, Configuration V ,
rises from a penalty of almost $6 million in 1972 to a saving of $26.22
million in 1980. In addition, $25.66 million of this saving is retained
with the elimination of depot level repair, making it very nearly expend¬
able.

All blade design characteristics have some impact on aircraft cosi
effectiveness. Of these, the most important are:

1. Blade scheduled retirement life.

2. Blade mean time between inherent damage.

3. Blade repairability.

4. Blade acquisition cost.

Significant changes in cost effectiveness are caused by variations in
blade scheduled retirement life, blade mean time between inherent
damage below 2,000 flight hours, and blade repairability. But blade
acquisition cost has, by far, the greatest impact on cost effectiveness.
This sensitivity is illustrated in Figures 52 and 53 for the baseline and
the most cost effective blade configurations for the 1972 and 1980 time
period.

125

•'t

co’

$ hS9^\/S30U>J - UOX
SS9U3AIJ39Jjg JSCQ IJEJOJiy

128

Figure 53. Impact of Blade Acquisition Cost,1980 Configurations.

These trends were obtained by varying blade acquisition cost in the cost-
effectiveness model while holding all other blade design characteristics
constant. The space between one configuration trend line and another is
due to differences in blade design parameters other than blade acquisition
cost. Blade repairability is the major contributor. Both Configurations
I and IV would be more cost effective if all blades had the same acquisi¬
tion cost. The inherent advantage of Configuration I over the baseline is
further enhanced by its lower acquisition cost for both time periods. The
relatively high acquisition cost of advanced technology Configuration IV
for 1972 overpowers its inherent repairability advantage, making it less
cost effective than the baseline. With the major reduction in blade ac¬
quisition cost available for 1980, this configuration becomes the most
cost effective 1980 design.

With the exception of blade acquisition cost, variations in blade design
characteristics do not produce blade expendability. Depot level expend-
ability is achieved when the elimination of depot level repair does not
incur a cost-effectiveness penalty (Figures 52 and 53). To achieve this,
the acquisition cost of the baseline blade would have to be reduced to
about $1, 700 for 1972 and $2,200 for 1980. Configurations I and IV would
have to be reduced to about $1,500 for either time period. Since these
derivative trends assume that all other design characteristics remain
constant, the expendability break-even points become invalid if design
changes made to reduce blade acquisition also change those other
characteristics.

Tables XX through XXIX present cost-effectiveness summary tables for
each configuration, followed by a discussion of each configuration's key
cost-effectiveness parameters. All of the candidate configurations are a
few pounds heavier than the baseline and yield a slight decrease in air¬
craft mission effectiveness. The greatest impact on cost effectiveness
derives from blade life-cycle cost. The prime contributor to this cost is
blade acquisition cost since it impacts directly on blade contributions to
flyaway cost, initial spares cost, and replenishment spares cost. Blade
mean time between inherent damage provides damage rate or the number
of blades damaged over an aircraft life cycle. Blade repairability deter¬
mines the number of damaged blades that are repaired and scrapped at
the various maintenance levels. An increase in blade repairability or
blade scheduled retirement life decreases the number of replenishment
spares required and therefore, decreases blade replenishment cost. All
of the candidate blade designs require more consumable tooling than the
baseline. This contribution to replenishment GSE cost is treated as an
increment over the baseline configuration, which is assumed at zero.

Tables XXX and XXXI summarize the blade life-cycle costs and the cost
of the new blade to the Army.

129

TABLE XX. COST EFFECTIVENESS SUMMARY,
BASELINE CONFIGURATION - 1972

Aircraft Mission Effectiveness 37.466 ton-knots

Aircraft Life-Cycle Cost $ 1,585,000

Aircraft Cost Effectiveness 23.638 ton-knots/megadollar

Fleet Effective Cost $ 1,585.00 megadollar

Fleet size adjusted to maintain
fleet effectiveness of 1,000 base¬
line UH-1 aircraft

Life Cycle Fuel and Oil Cost $ 53, 344

Blade Contribution To:

Flyaway cost

$ 6,000

Initial spares cost

$ 1,998

Replenishment spares cost

$ 36,202

Organizational level maintenance cost

$ 634

Direct support level maintenance cost

$ 554

Depot level maintenance cost

$ 3,071

Replenishment GSE cost

$ 0

Blade Life-Cycle Cost

$ 48,459

Life-Cycle Blades:

Damaged

10.94 Blades

Repaired at the organizational level

0 "

Repaired at the direct support level

1.31 "

Repaired at the depot level

2.03 "

Retired on schedule

0.81 "

Replenished by new spares

11.40 "

Expendability :

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23.601 ton-knots/megadollar

Fleet effective cost 1587.47 megadollar

TABLE XXI. COST EFFECTIVENESS SUMMARY,
_BASELINE CONFIGURATION - 1980

Aircraft Mission Effectiveness

37.466

ton-knots

Aircraft Life-Cycle Cost

$ 1,603,200

Aircraft Cost Effectiveness

23.370

ton-knots/mega -
dollar

Fleet Effective Cost

$ 1,603.20

megadollar

Fleet size adjusted to maintain
fleet effectiveness of 1,000 base¬
line UH-1 aircraft

Life-Cycle Fuel and Oil Cost

$ 53,34 4

Blade Contribution To:

Flyaway cost

8,417

Initial spares cost

2,771

Replenishment spares cost

49,979

Organizational level maintenance cost $

951

Direct support level maintenance cost $

608

Depot level maintenance cost

3,930

Replenishment GSE cost

Blade Life-Cycle Cost

66,656

Life-Cycle Blades:

Damaged

10.94 Blades

Repaired at the organizational level

0 "

Repaired at the direct support level

1.31 "

Repaired at the depot level

2.03 "

Retired on schedule

0.81 "

Replenished by new spares

11.40 ”

Expendability

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23.313 ton-knots/megadollar

Fleet effective cost 1607.08 megadollar

131

TABLE XXII. COST EFFECTIVENESS SUMMARY,
CONFIGURATION! - 1972

Aircraft Mission Effectiveness
Aircraft Life-Cycle Cost
Aircraft Cost Effectiveness

Fleet Effective Cost
Fleet size adjusted to maintain
fleet effectiveness of 1,000 base¬
line UH-1 aircraft

Life-Cycle Fuel and Oil Cost

37.437 ton-knots
$1,571,690

23.8 20 ton -knots/megadollar
$1,572.91 megadollar

53, 334

Blade Contribution To :

Flyaway cost
Initial spares cost
Replenishment spares cost
Organizational level maintenance cost
Direct support level maintenance cost
Depot level maintenance cost
Replenishment GSE cost

Blade Life-Cycle Cost

Life-Cycle Blades:

Damaged

Repaired at the organizational level
Repaired at the direct support level
Repaired at the depot level
Retired on schedule
Replenished by new spares

4,969

1,442

25,886

693

260

569

1,336

35,155

11.99 Blades
0.72
5.20 ”

0.47
1.13 "

9.73 "

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23.813 ton-knots/megadollar

Fleet effective cost 1573.36 megadollar

132

TABLE XXIII. COST EFFECTIVENESS SUMMARY,
CONFIGURATION 1-1980

Aircraft Mission Effectiveness
Aircraft Life-Cycle Cost
Aircraft Cost Effectiveness

Fleet Effective Cost
Fleet size adjusted to maintain
fleet effectiveness of 1,000 base¬
line UH-1 aircraft

37.437 ton-knots
$1,585,160

23.617 ton -knots /megadollar
$1,586.39 megadollar

Life-Cycle Fuel and Oil Cost

$ 53, 384

Blade Contribution To:

F lyaway cost
Initial spares cost
Replenishment spares cost
Organizational level maintenance cost
Direct support level maintenance cost
Depot level maintenance cost
Replenishment GSE cost

Blade Life-Cycle Cost

Life-Cycle Blades:

Damaged

Repaired at the organizational level
Repaired at the direct support level
Repaired at the depot level
Retired on schedule
Replenished by new spares

6,999

2,003

35,765

1,039

348

734

1,737

48,625

11.99 Blades
0.72 "

5.20 "

0.47 "

1.13 "

9.73 "

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23.607 ton-knots/megadollar
Fleet effective cost 1587.09 megadollar

133

134

135

TABLE XXVII. COST EFFECTIVENESS SUMMARY,
_CONFIGURATION III - 1980_

Aircraft Misrion Effectiveness
Aircraft Life-Cycle Cost
Aircraft Cost Effectiveness

Fleet Effective Cost
Fleet size adjusted to maintain
fleet effectiveness of 1,000 base¬
line UH-1 aircraft

Life-Cycle Fuel and Oil Cost

37.451
$1,593, 270
23.506

ton-knots

ton -knot s /megadollar

$1,593.91 megadollar

$ 53,339

Blade Contribution To:

Flyaway cost
Initial spares cost
Replenishment spares cost
Organizational level maintenance cost
Direct support level maintenance cost
Depot level maintenance cost
Replenishment GSE cost

Blade Life-Cycle Cost

Life-Cycle Blades:

Damaged

Repaired at the organizational level
Repaired at the direct support level
Repaired at the depot level
Retired on schedule
Replenished by new spares

7,646

2,303

41,337

980

371

2,750

1,342

56,729

11.85 Blades
0.24 "

2.58 ”

1.75 "

0.06 "
10.34 "

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23.442 ton-knots/megadollrr

Fleet effective cost 1598.23 megadollar

137

TABLE XXVIII. GOST EFFECTIVENESS SUMMARY,
CONFIGURATION IV- 1972

Aircraft Mission Effectiveness
Aircraft Life-Cycle Cost
Aircraft Cost Effectiveness

Fleet Effective Cost
Fleet size adjusted to maintain
fleet effectiveness of 1,000 base¬
line UH-1 aircraft

Life-Cycle Fuel and Oil Cost

37.431 ton-knots
$1,589,450

23.550 ton-knots/megadollar
$1, 590.92 megadollar

53, 336

Blade Contribution To :

Flyaway cost
Initial spares cost
Replenishment spares cost
Organizational level maintenance cost
Direct support level maintenance cost
Depot level maintenance cost
Replenishment GSE cost

Blade Life-Cycle Cost

Life-Cycle Blades:

Damaged

Repaired at the organizational level
Repaired at the direct support level
Repaired at the depot level
Retired on schedule
Replenished by new spares

9,939

1,985

36,613

922

1,306

490

1,660

52,915

16.59 Blades
0.15 "

12.16 ”
0.39 "

0.23 "

7.12 "

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23.529

Fleet effective cost 1592.35

ton-knots/megadollar
megadollar

TABLE XXIX. COST EFFECTIVENESS SUMMARY,
CONFIGURATION V - 1980

Aircraft Mission Effectiveness

37.431

ton-knots

Aircraft Life-Cycle Cost

$1,575,520

Aircraft Cost Effectiveness

23.758

ton -knots/megadollar

Fleet Effective Cost

$1,576.98

megadollar

Fleet size adjusted to maintain
fleet effectiveness of 1,000 base¬
line UH-1 aircraft

Life-Cycle Fuel and Oil Cost $ 53, 336

Blade Contribution To:

Flyaway cost

6,668

Initial spares cost

1,402

Replenishment spares cost

24,973

Organizational level maintenance cost

1,382

Direct support level maintenance cost

1,773

Depot level maintenance cost

631

Replenishment GSE cost

2,158

Blade Life-Cycle Cost $ 38,987

Life-Cycle Blades:

Damaged

Repaired at the organizational level
Repaired at the direct support level
Repaired at che depot level
Retired on schedule
Replenished by new spares

Expendability :

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23.748 ton-knots/megadollar

Fleet effective cost 1577.64 megadollar

16.59 Blades
0.15 "

12.16 "
0.39 "

0.23
7.12 "

139

TABLE XXX. COST EFFECTIVENESS

SUMMARY

Blade Design

Blade Life-Cycle Cost/Aircraft

1972

1980

....- -

Baseline UH-i

$48,459 (1)
$50, 925 (2)

$66,656 (1)
$70,540 (2)

Configuration I

$35,155 (1)
$35, 607 (2)

$48,625 (1)
$49,320 (2)

Configuration II

$36,807 (1)
$37,307 (2)

$51,106 (1)
$51,877 (2)

Configuration III

$66,435 (1)
$72, 922 (2)

$56,729 (1)
$61,050 (2)

Configuration

$52,915 (1)
$54,340 (2)

Configuration V

$38,987 (1)
$39,649 (2)

(1) Includes depot repair

(2) Eliminates depot repair

140

SIGNIFICANT CONFIGURATION FEATURES

Baseline Configuration

Mission effectiveness is established for the baseline aircraft using the
mission analysis program. The $3,000 Bell blade acquisition cost im¬
pacts directly on flyaway cost, initial spares cost, and replenishment
spares cost. This configuration's low repairability results in high blade
scrappage and consequently, high blade spares requirements. The li¬
mited field repairability, in particular, places a large burden on depot
level blade maintenance. 1980 labor rates and material costs increase
all blade life-cycle cost contributions and further decrease cost effective¬
ness.

Configuration I

This configuration has greater repairability, particularly on a field level,
which significantly reduces spares requirements. Blade acquisition cost
is less ihc.i for the baseline, directly reducing flyaway cost. Taken
with the reduced spares requirement, it provides a major reduction in
spares cost. Even with an increase in replenishment GSE cost, this
overall life-cycle cost reduction provides the aircraft with a major
increase in cost effectiveness. All blade life-cycle costs increase for
the 1980 time period , but the advantage over the baseline blade is re¬
tained. This is the most cost effective blade design for the 1972 lime
period.

Configuration II

This configuration is more cost effective than the baseline for both 1972
and 1980. Its damage rate and repairability are roughly comparable to
Configuration I, but it has a higher blade acquisition cost and hence,
yields a slightly lower aircraft cost effectiveness.

Configuration III

The repairability of this configuration is better than the baseline but
poorer than Configurations I and II. The damage rate is roughly compa¬
rable to Configurations 1 and II; but with less repairability, blade spares
requirements fall between these configurations and the baseline. For
1972, the high blade acquisition cost generated by the use of advanced
technology materials make;; this configuration less cost effective than the
baseline. For 1980, the lower blade acquisition cost is refected in a
marked reduction in blade life-cycle cost. For this time period, this
blade design becomes more cost effective than the baseline.

142

Configuration IV (1972) and V (1980)

This configuration exhibits the highest damage rate and the best repair -
ability of all blade configurations. Blade repairability is so superior,
however, that the combined effect produces the lowest spares require¬
ment. Repair costs remain moderate since 97% of repairs occur on the
field maintenance level. The use of expensive advanced technology ma¬
terial results in the largest 1972 blade acquisition cost of any configura¬
tion with a corresponding blade life-cycle cost penalty. For the 1972
time period, this penalty reduces cost effectiveness below that of Con¬
figurations I, II, and the baseline. The better blade repairability pro¬
vides a cost effectiveness advantage over Configuration III. With its low
spares requirement, this configuration is the most sensitive to reduction
in blade acquisition cost. With the application of the lower 1980 blade
acquisition cost, blade life-cycle cost and aircraft life-cycle cost are
greatly reduced, making this the most cost effective 1980 blade configura¬
tion.

Configuration VI

This configuration is simply an extension of Configuration I having all the
same features at the basic design. The automated pultrusion process has
been added for the trailing edge and should reduce manufacturing cost. A
complete analysis was not performed for this configuration. However,
it is felt that the pultrusion process could be productionized for this con¬
figuration by 1975 if development were to start by 1972.

143

DESIGN SELECTION

RATIONALE

Selection of the most cost effective rotor blade design from the four can¬
didate concepts discussed in this report was difficult since each design
offered some improvements over the present UH-1 blade. These im¬
provements were in the areas of reduced acquisition cost, improved re-
pairability, potential for highly automated production, and improved
rotor system performance. Design Configuration V, which is the twin
beam composite blade with the truss-type trailing edge, was finally
chosen as the prime candidate for 1980 based on the cost effectiveness
studies shown in Table XVIII and Figure 54 and by a comparative evalu¬
ation of the various blade attributes.

The acquisition costs of the four designs under consideration were com¬
pared for 1972 and 1980. The aluminum extruded spar with a fiberglass
trailing edge (Configuration I) had the lowest cost for 1972. The roll-
formed sheet metal spar and fiberglass trailing edge (Configuration II)
was slightly more costly than the aluminum spar design. Even though the
production of the spar channel sections of stainless steel and aluminum of
Configuration II was highly automated, it was more costly to bond and
assemble than the two-piece aluminum extrusion of Configuration I. The
twin beam spar, Configuration IV, and the "D" shaped tubular spar com¬
posite, Configuration III, are both considerably more expensive in 1972
than the metal spar blades because of the high cost of carbon and the need
for high automation.

The costs of all four candidate designs were projected for the P80 time
frame and are shown in Figure 54 along with the 1972 blade costs. The
figure shows that both metal spar blades are expected to increase sub¬
stantially in cost by 1980. Both the material and labor rate costs are ex¬
pected to increase as shown by Figure 55, while the fabrication time will
remain about the same because there is little room for increasing the
manufacturing technology (automation) for these metals. The composite
blades, however, have vast areas for potential improvements in manu¬
facturing . The methods to extrude a composite half section of a blade
will be developed in the next several years, a process which will save on
labor and also reduce the amount of wasted material common to present
composite production. This type of operation would also utilize low-cost
forms of the raw composite materials, such as spool roving, mat and
liquid resin, thereby reducing the costs further. In addition, the cost of
the composite materials, in particular carbon fiber, is expected to re¬
duce substantially in the next few years. Increased use of composite
materials, coupled with improvements in the manufacturing methods for
the basic materials is expected to reduce carbon composite prices to

144

Labor Rate Trends

Material Rate Trends

Figure 55.

Forecast of Material and Labor Costs.

$25. 00 per pound. Such reductions in material and labor costs are ex¬
pected to reduce the composite blade prices below those of the metal
spar designs by 1980.

Of the two composite blade designs, the twin beam spar concept will offer
the greatest potential for automated production because of its unique
simple two-half construction. The solid cross section of the spar with
its very simple, almost rectangular shape will be very easy to extrude.
The trailing edge truss section will also be extrudable, after develop¬
ment, resulting in an entire half section of the blade being fabricated as
one part. Making the blade in two halves eliminates the operations of
fabricating and assembling a separate skin, spar, spline, etc., for each
half section. The two halves are twisted, requiring very little torque,
and are assembled in separate precision-made contoured molds having
the blade twist. The blade thickness tolerance buildup is eliminated when
the two halves are machined off to a flat mating surface on the chordline
joint face and then bonded into one assembly.

The repairability of each design was also considered in the selection of
the prime candidate. Both of the metal spar designs have highly repair¬
able fiberglass trailing edge pockets. The fiberglass material has a low
notch sensitivity and can therefore tolerate a large amount of damage and
repair. The metal spars can tolerate only very minor damage and re¬
pair since the metals are much more notch sensitive. Projectile type
damage which punctures the metal spar would not be repairable because
of the high stress concentrations produced around the discontinuity in the
structure. Strain allowables of the metals used for the spar are consid¬
erably lower than the fiberglass strain allowables, which means that at
a given blade loading condition, the fiberglass will be operating at a much
higher margin of safety than the metals. Since the margins of safety of
the fiberglass components are higher, it follows that they will be more
tolerant of repaired damage than a similar metal part.

Repairability of the composite blades will be better than for either of the
metal spar designs because some repair of the spars is possible. Trail¬
ing edge repairability of the all composite blades will be about the same
as for the metal spar blades, since the trailing edge construction of all
the designs is very similar. The fiberglass spars both have t high mar¬
gin of safety and low notch sensitivity, making repair of even projectile
type damage possible. The twin beam spars are more repairable than
the tubular type spar because of the very simple configuration. Damaged
portions of the spar can be routed out and replaced with a bonded-in pre¬
molded repair section. The damaged honeycomb core would be filled
with a room temperature curing foam. The repair procedures are des¬
cribed in Appendix II.

147

Evaluation of the growth potential of the four candidate designs considered
the capability of each design to be used at higher aircraft speeds and also
to increase hovering performance by the increasing of blade twist. In¬
creases in blade twist will produce improved hover performance as
described in Appendix I, but will also increase blade vibratory stress
levels. Both the aluminum extruded spar and the roll-formed stainless
steel and aluminum spars are operating at close to their vibratory stress
limits. Very little increase in aircraft speed or blade twist can be toler¬
ated in either of these designs; therefore, there is little growth potential.
The composite blades, because of their much higher strain allowable
materials, have a very large potential for future growth. Either blade
twist or aircraft speed can be substantially increased, or a combination
ot both, without overstressing the blade.

148

CONCLUSIONS

The following conclusions are based upon a study of over 15 blade designs
which are interchangeable with the UH-1 blade. The conclusions pertain to
structural skin designs which are limited to the UH-1 requirement for
edgewise rigidity. The conclusions could be quite different for blades
with nonstructural pockets which are used extensively in articulated
rotors.

A. 1972 Time Frame - Configuration I

The aluminum spar blade with a fiberglass and honeycomb cover is
the optimum expendable blade configuration for the 197 2 time
frame. This blade requires very little development and could be
retrofitted immediately. The blade has 30% fewer parts than the
Bell blade. It uses 6061 aluminum spars instead of 2024 aluminum
for superior corrosion characteristics. The spar has a 1. 3-inch-
thick leading edge for erosion protection. The blade is 54% repair¬
able compared with 31% for the Bell blade. Its life-cycle costs
could result in a $12 million savings. The steel blade was slightly
more expensive,and the composite blade material costs in this time
frame were prohibitive.

B. 1975 Time Frame - Configuration VI

An aluminum spar with an automated advanced composite cover was
considered to be the optimum expendable blade for this time frame.
Essentially, this is a composite blade with an aluminum spar. Be¬
cause 70% of the blade costs are associated with the cover assembly,
the use of the pultrusion method of manufacturing a one-piece fiber¬
glass cover offers tremendous savings potential. It is considered
that this technology can be demonstrated by 1975. The aluminum
spar is the same as that defined in Section A.

C. 1980 Time Frame - Configuration V

The all composite Sikorsky twin beam design can potentially be the
optimum expendable blade for the 1980 time frame. For this po¬
tential to be realized, the cost of carbon or boron must be reduced
to $25 a pound. The ability to withstand damage without shattering
or delaminating extensively must be demonstrated. The ability to
repair the structural spar and trailing edges without significantly
affecting its strength must be demonstrated. Finally, the pultru¬
sion or other automated methods of fabricating the blade in one or

149

two pieces with contour and weight control must be demonstrated.
It is considered that this technology must be proven prior to com¬
mitting to production. There is sufficient time to develop this
technology, and the twin beam concept should greatly simplify the
development task.

150

RECOMMENDATIONS

It is recommended that the pultrusion process of manufacturing a one-
piece composite cover be developed. It is applicable to the aluminum
spar concept for 1975 and the twin beam composite blade concept for
1980. In addition, it is recommended that a program be undertaken to
develop the twin beam concept and demonstrate the needed technology as
early as possible. It is conceivable that this concept could be available
much sooner than 1980. To explore this, it is recommended that the
twin beam design be manufactured, fatigue tested, whirled and flown.

In parallel, the pultrusion development should be expanded to include the
entire blade.

151

LITERATURE CITED

1. DESIGN STUDY OF REPAIRABLE MAIN ROTOR BLADES, Kaman
Aerospace Report R-928, April 1971.

2. Carr, P. V., and Hensley, 0. L., UH-1 and AH-1 HELICOPTER MAIN
ROTOR BLADE FAILURE AND SCRAP RATE DATA ANALYSIS, Bell
Helicopter Company, USAAVLABS Technical Report 71-9, Eustis
Directorate, U. S. Army Air Mobility Research and Development
Laboratory, Fort Eustis, Virginia, January 1971, AD 881132L.

3. ORGANIZATIONAL MAINTENANCE MANUAL - ARMY UH-1D HELI¬
COPTER, Army TM-55-1520-210-20.

4. DIRECT SUPPORT AND DEPOT MAINTENANCE MANUAL, Army
TM-55-1520-35.

5. Carlson, R. G. , and Hilzinger, K. D. , ANALYSIS AND CORRE¬
LATION OF HELICOPTER ROTOR BLADE RESPONSE IN A VARI¬
ABLE INFLOW ENVIRONMENT, Sikorsky Aircraft Div. , USAAML
Technical Report 65-51, U. S. Army Aviation Materiel Laboratories,
Fort Eustis, Virginia, 1965, AD 622412.

6. MIL-HDBK-5 METALLIC MATERIAL AND ELEMENTS FOR AERO¬
SPACE VEHICLE STRUCTURES, Department of Defense, Washing¬
ton, D. C.

7. Bell, W. J. , and Benham, P. P. , THE EFFECT OF MEAN STRESS ON
FATIGUE STRENGTH OF PLAIN AND NOTCHED STAINLESS STEEL
SHEET IN THE RANGE FROM 10 TO 10 8 CYCLES, Symposium On
Fatigue Tests of Aircraft Structures, ASTM STP 338, 1963, pp 25 -
46.

8. MIL-HDBK-17A PLASTICS FOR AEROSPACE VEHICLES, PART I,
REINFORCED PLASTICS, Department of Defense, Washington, D. C.,
January 1971.

9. STATIC AND FATIGUE TEST PROPERTIES FOR WOVEN AND NON-
WOVEN S-GLASS FIBERS, Boeing Vertol Co., USAAVLABS Techni¬
cal Report 69-9, U. S. Army Aviation Materiel Laboratories, Fort
Eustis, Virginia, April 1969, AD 688971.

10. Roark, Raymond J., FORMULAS FOR STRESS AND STRAIN, New
York, New York, McGraw-Hill Book Co. , Inc., 1954.

152

11. Miner, M. A., CUMULATIVE DAMAGE IN FATIGUE, Journal of
Applied Mechanics, 1945.

12. CHARACTERIZATION OF BORON, GRAPHITE AND GLASS FILA¬
MENT/ORGANIC MATRIX COMPOSITE MATERIALS, Sikorsky
Aircraft Engineering Report SER-50644, January 1970.

13. MATERIALS PROPERTY HANDBOOK VOL. II STEELS, AGARD
(1/2 Hard 150/110),March 1966.

14. ENGINEERING MATERIALS MANUAL, Sikorsky Aircraft,1972.

15. UH-ID HORSEPOWER REQUIREMENTS STUDY, CORG Memo 185,
June 1965.

16. COMBAT OPERATIONAL FLIGHT PROFILES ON THE UH-1G,
AH-1G, AND UH-1H HELICOPTERS, AHS Forum, June 1970.

17. FY 71 DETAIL SPECIFICATION 205-947-135, March 2, 1970.

18. CATEGORY II PERFORMANCE TESTING OF THE YUH-1D WITH A
48 FOOT ROTOR, FTC-TDR-64-27.

153

APPENDIX I

BLADE CHARACTERISTICS

EFFECT OF TORSIONAL STIFFNESS ON BLADE STRESS AND TOR-

StONAL DEgLECnSH -

A study was made to determine the effect of varying the torsional rigidity
on various blade parameters. The UH-1H rotor blade was used as the
base for the study. Torsional stiffness of the UH-1H blade was both in¬
creased and decreased to observe the effect on blade stress and blade
torsional deflection. The blade stiffness was varied by the same per¬
centage from root to tip of the blade in each case studied. It was ob¬
served that the torsional deflection and blade stress were directly re¬
lated to each other; as deflection began to rise rapidly when the stiffness
fell to approximately 50% of the base value, the stress also began to rise
rapidly. The relationship between torsional deflection in degrees and
the percentage of the UH-1H torsional stiffness is shown in Figure 56.
Vibratory stress normalized to the UH-1H stress at the UH-1H stiffness
is plotted vs. the varying UH-1H torsional stiffness in Figure 57. In all
cases, the study was conducted at a forward velocity of 110 knots and a
gross weight of 3500 pounds. The study has shown that reduction in tor¬
sional stiffness of up to 50% of the UH-1H is possible without serious
problems; however, for a blade with increased forward speed potential,
the torsional stiffness must be kept at a high level.

EFFECT OF BLADE TWIST ON HOVER PERFORMANCE

A study was made to determine the improvements which can be made in
rotor hovering performance by varying the blade twist. The Sikorsky
Rotor Hover Performance Analysis which is programmed for a UNIVAC
computer was used for the calculations. The UH-1H rotor system was
used as the model for these calculations. Because the basic UH-1H blade
loading is not very high (Cy/<f =. 0854), the performance gains were
relatively small. About 50 pounds of additional thrust was obtained for
each additional degree of blade twist at 990 horsepower delivered to the
rotor. Figure 58 illustrates the improvement in rotor thrust vs blade
twist.

154

156

Figure 58. Rotor Thrust - Blade Twist Curve.

APPENDIX II

RELIABILITY/MAINTAINABILITY DATA

Appendix II includes the Reliability/Maintainability data referred to in
the body of this report. It consists of Configurations I, II, and IV
Reliability Analyses, Repairability Analyses, Math Model R/M Input
Vauables, Failure Mode and Effects Analysis and Repair Procedures.

158

TABLE XXXII. RELIABILITY ANALYSIS - CONFIGURATION

00 _
<L> ®

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L59

15.0

TABLE XXXII. (Continued)

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General Overstressed, warped, sudden stop, overspeed 126.0

Total External 826. 0
Total Blade 1201. 0

TABLE XXXIII. (Continued)

OOOqOOOOoO

CNCS^f-Iodt^O^cC ^t*
to r—t i—i i—i oo co 'sC' co

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165

14. 0

166

TABLE XXXIV. MATH MODEL R/M INPUT VARIABLES -

«***-«* WW'Of

167

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L68

Water contamination 2. Vibration due to blade unbalance. 2. Blade unbalance. leaVage or seepage

169

170

171

172

174

175

179

TABLE XXXVIII. MATH MODEL R/M INPUT VARIABLES - CONFIGURATION II
Variable Value

180

182

183

upon degree and location.

TABLE XXXIX. (Continued)

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184

186

TABLE XXXX. ( Continued )
Externally Caused

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188

189

TABLE XXXXI. (Continued)

190

TABLE XXXXI. (Continued)

Msis-

Dented 22.0 87.0

Foreign object damage 31.0 42.0

TABLE XXXXI. (Continued)

TABLE XXXXtI. MATH MODEL R/M INPUT VARIABLES -

CONFIGURATION IV

Variable Value

L. Aircraft Down Hours

2. Aircraft Aborting Failure Rate

3. Blade Mean Time Between In¬
herent Damage

4. Blade Retirement Life

5. % Damage Repaired On Aircraft,
ORG Level

6. % Damage Repaired Off Air¬
craft,ORG Level

7. % Removed Blades Scrapped at

ORG Level

8. % Removed Blades Sent to Direct
Support

9. % Received Blades Repaired at
Direct Support

0. % Received Blades Scrapped at
Direct Support

1. % Received Blades Repaired at
Depot

2. Maintenance Man-Hours to In¬
spect On Aircraft (ORG)

3. Maintenance Man-Hours to Re¬
pair On Aircraft (ORG)

4. Maintenance Man-Hours to Re¬
pair Off Aircraft (ORG)

5. Maintenance Man-Hours per
Blade Repair (Direct Support)

6. GSE Cost Per Repair (Direct
Support)

7. GSE Cost Per Aircraft (Direct
Support)

8. Parts/Material Cost (Direct
Support)

9. Blade Overhaul Cost (Depot)

4. 3839 Down hours per flight
hour

. 015 Aborting failures per
flight hour

1, 200 Blade hours

5, 000 Blade hours

0.9 Percent
0 Percent
12. 0 Percent

87.0 Percent

85.0 Percent

10.0 Percent

55. 0 Percent

. 25 Maintenance man-hours

l. 0 Maintenance man-hours

7.5 Maintenance man-hours

$84.10 per repair

$637. 66 per aircraft

$66. 00 per repair
$822.00 per blade

193

194

195

196

Possible Failure Modes in Effect of Failure Upon Assembly Probable Symptom and

Nomenclature Function Anticipated Knvironment Function Method of Detection

’vy*-w' ■»-

197

iuvuep jo

REPAIR PROCEDURES

1. LEADING EDGE POLYURETHANE RESTORATION

1. Clean and strip area to be restored using a rag saturated with
methyl ethyl ketone.

2. Fill porous laminate substrates with a patching paste or filler.

3. Apply primer coat in accordance with instructions supplied with
kit. Allow primer to dry at room temperature for not less than
1 hour. Primer coat may be sanded with 180 or 220 grit paper
to promote adhesion and to provide smooth finish.

4. Mix and apply basecoat of polyurethane vehicle in accordance with
instructions. Apply coating by brush or spray gun to achieve a dry
mil thickness of 2 mils per coat.

5. Apply at 30 minute intervals to obtain a recommended basecoat
minimum thickness of 6 mils.

6. Mix and apply final coat in accordance with instructions. The fi¬
nal coat should be applied with brush or spray gun to achieve a dry
mil thickness of 2 mils per coat.

7. Apply at 30 minute intervals to obtain a recommended final coat
thickness of 14 mils.

8. Allow to cure for minimum of 24 hours at 70 - 75°.

198

2. LEADING EDGE CARBON REPAIR

Foreign object damage to the leading edge not exceeding 2 inches in span-
wise length and not extensive enough to have penetrated to the twin beam
and honeycomb core is repairable. The damaged area would be replaced
with a 6-inch-long prefabricated prescarfed section.

1. Using the prefabricated section as a template, clean an overlap area
on the blade 4 inches larger than the prefabricated section.

2. Cut out the damaged area of the blade to allow the prefabricated
section to be fitted into place.

3. Using the scarfing tool provided, scarf the skin adjacent to the cut¬
out area for mating with the prefabricated replacement.

4. Cut an overlay patch from the skin material 2 inches larger than
the scarfed area.

5. Clean the repair area, the prefabricated replacement and the over¬
lay patch by wiping with methyl-ethyl-ketone and applying primer.

6. Prepare adhesive and apply to the repair area and the inside of the
prefabricated replacement section.

7 Positior -eplacement section .

8. Apply adhesive to the inside surface of the overlay patch and to the
outside of the repair area.

9. Position overlay patch and tape with masking tape.

10. Cover patch area with wax paper or Teflon sheet and position com¬
pression blanket over entire area. Inflate compression blanket to
15 PSI and leave in place for 8-hour cure period.

11. Remove compression blanket and finish patch by sanding edges to
remove excess material and achieve a feathered edge.

199

3. LEADING EDGE BLEND REPAIR

A hollow aluminum alloy extrusion forms the leading edge of the blade.
Nicks, dents or gouges in this aluminum leading edge or spar would be
repairable within certain limits established for various areas of the
spar. As an example, damage to the leading edge up to .250-tnch deep
would be repairable. All repairs would be made in accordance with the
following instructions;

1. From the repair limitations figure, determine the maximum
depth of damage that maj be repaired in that area.

2. If the damage is repairable, use a single cut mill file to
remove the damage, limiting the cleaning strokes to a span-
wise motion. Do not file or blend in a chordwise direction.

3. Remove all file marks and blend the area, using #150 aluminum
oxide abrasive cloth so that the depth of the repair is at least
0. 002 inch deeper than the depth of the damage, but no deeper
than the limits established for that area.

4. The width of the blend area must be 30 times the depth of the
final rework.

5. Apply Alodine 1200 solution, specification MIL-C-5541, to the
repaired area. Allow to remain on the surface for 2 to 4
minutes, wash with water and wipe dry.

200

4. TWIN BEAM REPAIR

The following repair procedure can be accomplished on the unidirectional
fiberglass internal beams that have been damaged by a foreign object
that penetrated the carbon skin.

The required repair materials, supplied in kit form, would include taper¬
ed premolded unidirectional patches, skin materials, a router and a
router template.

1. Strip and clean the affected area and enough of the surrounding
skin to afford a 4-inch overlap. Allow a 2-inch overlap

for patches less than 1 square foot in area. Wipe down
with methyl-ethyl-ketone (MEK).

2. Remove the skin from the damaged area to explore enough of
the beam to allow routing and patching of the beam, by cutting
through the skin with a sharp tool and applying heat to loosen
the skin-to-beam and skin-to-core bond. Skin should be re¬
moved from a minimum of 4 inches around the damaged area.

3. Using a template and router, remove the damaged beam
material to accept a tapered premolded beam patch.

4. Clean the tapered repair area of the beam by wiping with
methyl-ethyl-ketone and applying primer.

5. Check the fit of the premolded patch to the repair area. Cut
patch to proper width if necessary.

6. Select appropriate skin patch and trim to fit flush into the skin
cut-out area.

7. Clean both sides of skin patch by wiping with methyl-ethyl-
ketone and prime both sides.

8. Cut overlay patch to size and prepare for installation by wiping
with methyl-ethyl-ketone and applying primer.

9. Prepare adhesive and apply first to the beam repair area and
the tapered premolded patch. Position patch flush in the beam
repair area.

10. Apply adhesive to the repaired beam and the skin patch; position
skin patch flush in the cut-out area.

201

11. Apply adhesive to the outside surface of the skin patch, the
surrounding overlap area and to the inside of the overlay patch.

12. Cover the patch with scrim cloth trimmed to the outer edge of
the patch.

13. Position overlay patch over the patched area. Allow equal
overlap on all sides.

14. Cover the entire patch with wax paper or Teflon sheet and
position compression blanket over entire area. Inflate com¬
pression blanket to 15 PSI and leave in place for 8-hour cure
period.

15. Remove compression blanket and finish patch by sanding edges
to remove excess material and achieve a feathered edge.

202

5. FIBERGLASS SKIN REPAIR - PATCH

The following repair procedures can be accomplished for fiberglass skin
damage such as abrasion, delamination, bonding separation, crazing and
blisters. The repair procedure is not limited by size of affected area and
has been successfully accomplished and flight tested at Sikorsky Aircraft.
This procedure, with variation, is also applicable to external damage
which penetrates the skin and extends into the honeycomb core. The re¬
pair procedure illustrated in Figures 59 through 72 was actually perform¬
ed on a recent flight test article to repair extensive skin to honeycomb
bonding separation which occurred due to damage to the trailing edge re¬
sulting from contact with a foreign object. All required repair materials
including skin patches up to 1 square foot can be furnished in kit form.

Skin material required for patches exceeding 1 square foot in area can be
supplied in bulk form.

1. Strip and clean the affected area and enough of the surrounding skin
to afford a 4-inch overlap. Allow a 2-inch overlap for patches less
than 1 square foot in area. Wipe down with methyl ethyl ketone
(M. E. K.).

2. Remove damaged skin by cutting around periphery of affected area
with a sharp tool and applying heat to loosen the skin-to-core bond.

3. Select appropriate patch and trim to fit flush into cut-out area.

4. Prime both sides of patch and skin overlap area.

5. Prepare adhesive and apply to small separations between skin and
core around periphery of cut-out area.

6. Apply adhesive to patch and position patch flush in cut-out area.

7. Apply adhesive to outside surface of patch and surrounding overlap
area.

8. Cut overlay patch to size and prepare for installation by wiping with
methyl ethyl ketone and applying primer.

9. Apply adhesive to surface of overlay patch and cover with scrim
cloth. Trim scrim cloth to outer edge of patch.

10. Position overlay patch over patched area. Allow equal overlap on
all sides.

203

11. Cover entire patch with wax paper or Teflon sheet and position com¬
pression blanket over entire area. Inflate compression blanket to
15 PSI and leave in place for 8-hour cure period.

12. Remove compression blanket and finish patch by sanding edges to
remove excess material and achieve a feathered edge.

204

,pply Adhesive & Position Patch. Figure 66. Apply Adhesive & Position Patch

Vv ■ \

1 I

■gfe.l-*. t :

wmam • ■!

V ' • •■ '-

Figure 71. Install and Inflate ComDression Blanket.

Figure 72. Finished Patch.

208

•»**w«Tv* *s»cvt*

6. FIBERGLASS SKIN REPAIR - PLUG

The following repair procedure can be accomplished when a foreign ob¬
ject penetrates the fiberglass skin and honeycomb. Prefabricated skin/
core sections of varying sizes would be furnished in kit form to replace
damaged sections. Figures 73 through 84 illustrate this type of repair.

1. Strip and clean the affected area and enough of the surrounding skin
to afford a 4-inch overlap. Allow a 2-inch overlap for patches
less than 1 square foot in area.

2. Remove damaged skin by cutting around periphery of affected area
with a sharp tool. Note: Wraparound template shall be used to
align top and bottom holes.

3. Force a hacksaw blade through honeycomb and cut hole through
blade, working alternately from top and bottom sides of blade.

4. Cut skin/core plug to fit hole.

5. Install skin/core plug and tape in position on top side.

6. Sand excess honeycomb on bottom side.

7. Remove plug, clean plug and hole with methyl-ethyl-ketone.

8. Cut two overlay patches to size and prepare for installation by
wiping with methyl-ethyl-ketone and applying primer.

9. Prepare adhesive and apply to plug and hole.

10. Install plug.

11. Apply prepared adhesive to top of plug, overlap area of skin,and to
one side of one patch.

12. Position overlay patch over patched area. Allow equal overlap on
all sides.

13. Tape patch in position and cover with waxed paper or Teflon sheet.

14. Repeat Steps 11 through 13 for bottom side.

15. Position compression blanket over entire area. Inflate compression
blanket to 15 PSI and leave in place for 8-hour cure period. Remove
blanket and sand edges of patch to blend into blade contour.

209

210

1 7"*-’*

• ’

•

ISI\ V r

Gi^A \

m ]

N&La \

-- » r.

;»

L./

- i .««

7 • ATTACHING POINT BUSHING REPLACEMENT

Removal and replacement of worn, scored, corroded, or damaged attach¬
ing point bushings can be accomplished in the field since the subject blade
will be manufactured with dual steel bushings (the outer steel bushing
being adhesively bonded to the blade).

1. Press out damaged internal bushing using tool provided by the manu¬
facturer.

2. Chill the inner steel replacement bushing in a solution of dry ice
and methyl-ethyl-ketone (M.E.K.) for 3 minutes or until bubbling
stops.

3. Wipe off replacement bushing and install into blade.

213

8- FIBERGLASS TRAILING EDGE SPLINE REPAIR

Delaminations and foreign object damage to the trailing edge spline are
repairable. The damaged section would be replaced with a prefabricated
prescarfed section.

1. Strip and clean the affected area and enough of the surrounding area
to afford a 2-inch overlap chordwise and 4-inch spanwise.

2. Using the prefabricated prescarfed section as a template, lay out
the area to be removed.

3. Using a hacksaw blade, cut out the affected area.

4. Fit the prefabricated section into the repair area.

5. Cut two overlay patches to size allowing the specified overlap.

6. Clean repair area of the trailing edge, prefabricated section and the
two overlap patches by wiping with methyl-ethyl-ketone and applying
primer.

7. Prepare adhesive and apply to top side of blade overlay area and one
side of one overlay patch.

8. Position overlay patch and tape with masking tape.

9. Apply adhesive to the prefabricated section and fit into repair area.

10. Apply adhesive to one side of the second overlay patch and to the
bottom side of the blade overlap area.

11. Position second overlay patch and tape with masking
tape.

12. Cover patch areas with wax paper or Teflon sheet and position com¬
pression blanket over entire area. Inflate compression blanket to
15 PSI and leave in place for 8-hour cure period.

13. Remove compression blanket and finish patch by sanding edges to
remove excess material and achieve a feathered edge.

214

APPENDIX III

QOST-EFFECTIVNESS MODEL

The cost-effectiveness model computes the cost effectiveness of the UH-1
aircraft equipped with any candidate rotor blade design. The model has
been programmed for the (JNIVAC 1108 computer and is described by the
following sections:

1. Input definition

2. Mission effectiveness analysis

3. Blade utilizr ion °nd logistics analysis

4. Blade life-cycle cost analysis

5. Aircraft cost-effectiveness analysis

6. Mission analysis

A step-by-step description and output definition are given for Sections 2
through 5. An asterisk * is used to denote multiplication to avoid the
ambiguous alphabetical symbol, x. Model input variables are parenthe¬
sized in the equations for further clarity.

215

1.1 GENERAL INPUT DEFINITIONS

Symbol

Description

Units

A b

Blade set attrition

sets/FH

Installed blades per aircraft

BLCC uh

Baseline UH-1 blade life-cycle cost

CB acq

Single blade acquisition cost

G cont

Blade container cost

G fuel

Fuel and oil cost per pound of fuel
consumed

$/lb

CGR ^gp

Replenishment GSE cost per repair,
depot level

CGR ds

Replenishment GSE cost per repair,
direct support level

°° R 0«

Replenishment GSE cost per off-
aircraft repair, organizational level

CGR o„

Replenishment GSE cost per on-
aircraft repair, organizational level

GGS dep

GSE support cost per aircraft,depot
level

008 d.

GSE support cost per aircraft, direct
support level

ccs o

GSE support cost per aircraft,
organizational level

C m

Average mission capability

ton-knots

CMR ds

Mean material cost per blade repair,
direct support level

CMR off

Mean material cost per off-aircraft
blade repair, organizational level

216

f ? 5 *» ’***¥‘2*1<,vx>-- ., ...

Symbol

CMR

dep

C^UH

CSH

cont

Description Units

Mean material cost per on-aircraft $

blade repair, organizational level

Blade overhaul cost, depot level $

Baseline UH-1 life-cycle fuel and oil $
cost

Empty blade container shipping cost $
from field to CONUS

cs Hfld

Packaged blade shipping cost from $

CONUS to field

CSHP

dep

CSHP

«%s

E mUH
FF

Blade shipping preparation cost, $

depot level

Blade shipping preparation cost, $

direct support level

Blade shipping preparation cost, $

organizational level

Packaged blade shipping cost from $

field to CONUS

Aircraft down hours per flight hour DH/FH

Baseline UH-1 mission effectiveness ton-knots

Average mission fuel flow lb/FH

K .
inv

Ratio of blade inventory spares to
blade life-cycle replenishment spares

LCC

inst

Aircraft service life FH

Blade scheduled retirement life FH

Baseline UH-l life-cycle cost $

Mean maintenance man-hours per blade MMH
installation

217

Symbol

Description

Units

MI.

dep

Mean maintenance man-hours per blade
receiving and inspection, depot level

MMH

MI J
ds

Mean maintenance man-hours per blade
inspection, direct support level

MMH

off

Mean maintenance man-hours per off-
aircraft blade inspection, organizational
level

MMH

MI on

Mean maintenance man-hours per on-
aircraft damaged blade inspection,
organizational level

MMH

M rem

Mean maintenance man-hours per blade
removal

MMH

req

Mean maintenance man-hours to
requisition and obtain a replacement
blade, organizational level

MMH

MRds

Mean maintenance man-hours per blade
repair, direct support level

MMH

MR off

Mean maintenance man-hours per off-
aircraft blade repair, organizational
level

MMH

MR on

Mean maintenance man-hours per on-
aircraft blade repair, organizational
level

MMH

dep

Mean maintenance man-hours per blade
scrappage, depot level

MMH

MS ds

Mean maintenance man-hours per blade
scrappage, direct support level

MMH

MS o

Mean maintenance man-hours per blade
scrappage, organizational level

MMH

MTB e

Blade mean time between external
damage

218

Symbol

Description

Units

MTB.

Blade mean time between inherent
damage

PB ds

Percent of damaged and removed
blades sent to direct support

PBR dep

Percent of received blades repaired
at depot level

PBR ds

Percent of received blades repaired
at direct support level

PBR off

Percent of damaged and removed
blades repaired at organizational level

PBR on

Percent of damaged blades repaired
on aircraft

PBS ds

Percent of received blades scrapped at
direct support level

p B s 0

Percent of damaged and removed blades
scrapped at organizational level

civ

Civilian maintenance personnel labor rate

$/hr

R mil

Military maintenance personnel labor rate

$ dir

Aircraft mission abort failures per flight
hour

maf/FH

T m

Average mission flight time

u a

Aircraft annual utilization

219

1.2 NONVARIABLE INPUTS

The following inputs described in Section 1.1 are assumed not to vary
with rotor blade design:

Customer Specified
input Value

Contractor Specified
Input Value

A b

.0003

blcc uh

48455

C fuel

.02

cont

200.

CP°L UH

53344

CSH

cont

45.

mUH

37. 466

CSH fid

130.

K .
inv

. 05263

CSHP ,
dep

70.

lcc uh

1,585,000,

CSHP ,
ds

70.

CSHP

70.

MS o

CSH US

90.

u a

500,

L a

5000.

M m st

3. 75

^dep

2. 5

rem

3. 75

req

MS dep

R 'civ

12.00

R mil

4. 00

220

2.0 AIRCRAFT MISSION EFFECTIVENESS

The mission effectiveness of a single aircraft is the product of its mis¬
sion availability, mission reliability, and mission capability.

2.1 Average daily utilization - FH/day

U d = < u a) (28)

"365

2.2 Average daily downtime -hr/day

T d = <DT)*U d

2.3 Mission availability

A m 24 - T c

2.4 Mission reliability

R m = e

<Rs> * (T m )

2.5 Mi s s ion effectivenes s

* (C m )

221

3. 0 BLADE UTILIZATION AND LOGISTICS

The computation of blade life-cycle cost must reflect the maintenance,
replenishment, inventory, and shipping burdens imposed by the rotor
blade design. This analysis establishes the blade requirements of a
single aircraft throughout its life cycle.

3. I Blades inherently damaged. Based on aircraft retirement life and
specified blade mean time between inherent damage.

BDi = (B) * (L a )

(33)

(MTB^

3. 2 Blades externally damaged. Based on aircraft retirement life and
blade mean time between external damage.

BD„ = <B) *(L a )

(•cm; i

(34)

3. 3 Total blades damaged.

BD = BD i + BD e

(35)

3. 4 Damaged blades repaired on aircraft. Based on a specified per¬
centage.

BR _ _ (PBR on ) * BD

on -Too-

(36)

3. 5 Damaged blades removed from aircraft. All damaged blades not
repaired on aircraft.

BD rem _ BD - BR on

(37)

3. 6 Removed blades repaired off aircraft, organizational level. Based
on a specified percentage.

BB off = * BD rem

- m -

(38)

3. 7 Removed blades scrapped, organizational level. Based on a
specified percentage.

_ (PBS 0 ) * BD r em
- IT50-

(39)

222

3.8 Removed blades sent to direct support. Based on a specified
percentage.

3 = (”ds) BD rem /40\

ds —ioo- ( }

3. 9 Damaged blades sent to depot from organizational level. Removed
blades not scrapped, repaired at organizational level, or sent to
direct support.

B ^dep - BD rem " BR off " BB o " *Ms (41)

3.10 Damaged blades repaired, direct support. Based on a specified
percentage of blades received.

BRh« = ( PBRds ) * Bds
100

(42)

3.11 Damaged blades scrapped, direct support. Based on a specified
percentage of blades received.

3S = (PBS ds )* Bds
as -TOO-

3.12 Damaged blades sent to depot from direct support. Blades
received at direct support not scrapped or repaired.

(43)

BP)B dep B ds " BR ds " BS ds

3.13 Total damaged blades sent to depot.

B dep “ B(\iep + BP ^dep

(44)

(45)

3.14 Damaged blades repaired, depot. Based on a specified percentage
of blades received.

BR Ho „ = < PBR dep> * B dep
-TOO-

v dep “ (46)

3.15 Damaged blades scrapped, depot. Received blades not repaired.

B ^dep = B dep ' B^dep (47)

3.16 Total damaged blades scrapped, all levels.

BS - BS 0 4- BS ds + BS dep

(48)

3. 17 Blades lost to attrition. Based on blade set attrition rate and
aircraft retirement life.

Batt = (B) * (Af)) * (La) (49)

223

3. 18

Blades lost to scrappage and attrition.

B sa BB + B att

(50)

3.19 Aircraft mean time between loss of blades to attrition - flight
hours.

MTBa = (Ktf~ (51)

3.20 Aircraft mean time between inherent or external blade damage -
flight hours.

MTB d = 1

I-

(mT%7

3. 21 Aircraft mean time between blade scrappage - flight hours.

-j-

(MTB e )

(52)

MTB q = MTB d * BD
S -BS-

(53)

3.22 Aircraft mean time between scrappage and attrition - flight hours.

1 - <54)

M^sa

~T~

WTK

3. 23 Aircraft mean time between scrappage, attrition, or blade retire¬
ment - flight hours. The time between scrappage or attrition
may vary considerably from the mean, allowing some blades to
reach their retirement lives. If blade retirement life exceeds air¬
craft retirement life, no blades are retired. If not, the following
probability integral formula is used to estimate mean time
including retirement:

-d b )

MTB sar = MTB_ a Ml - e MTB. a )

(55)

3.24 Blade replenishment spares. The sum of blades scrapped,
retired, and lost to attrition.

R , _ (S) * (L a )

repl MTB,

(56)

'sar

3. 25 Blades retired from service.

ret

= B

repl

- B

(57)

224

3. 26 Blades removed or installed. The sum of blades removed due to
damage and blades retired.

B f i = BD rem + B ret (58)

3. 27 Blades requisitioned from inventory. All removed blades not
repaired at the organizational level.

B req = B ri " BR 0 ff ( 59 )

3. 28 Initial blade spares. Inventory blades either on hand or in the

supply pipeline. Assumed to be proportional to life-cycle replen¬
ishment spares requirement.

B inv = <^inv) * B repl (60)

225

4. 0 BLADE LIFE-CYCLE COST

This analysis computes blade contributions to the life-cycle cost of a

single aircraft.

4. 1 Blade contribution to aircraft flyaway cost. Based on the acquisi¬
tion cost of installed blades.

Cfly = (B) * (C®acq) (61)

4. 2 Blade contribution to initial spares cost. The cost of blades and
containers in the spares inventory and shipping from CONUS to
field.

C isp = p^acq) + ( c cont) + ( CSH fld^j * B inv (62)

4. 3 Blade contribution to replenishment spares cost. The cost of
replenishment blades and shipping in recycled containers.

Crsp = [(CBacq) + <CSH tld ) + (CSH^jl * B repl (63 )

4. 4 Cost of on-aircraft inspection for blade repairability, organization¬
al level. Based on a specified mean MMH per damaged blade.

a on * < R mil> * <»«on> * BD (64)

4. 5 Cost of on-aircraft blade repairs, organizational level. Based on a
specified mean MMH and material cost per blade repair.

C^on = J(R m il) * (MR 0 n) + (CMR on )j * BR on (65)

4. 6 Cost of blade removal, organizational level. Based on mean MMH
per blade removal.

^rem = (^mil) * (M rem ) * B r | (55)

4. 7 Cost of off-aircraft inspection for blade disposition, organizational
level. Based on mean MMH per damaged blade removed.

CI 0 ff = (^mil) * (MIoff) * B ^rem (67)

4.8 Cost of off-aircraft repairs, organizational level. Based on a
specified mean MMH and material cost per blade repair.

CR 0 ff = [(Rmil) * (MR 0 ff) + (CMR off )j * BR off (68)

226

4. 9 Cost to requisition and obtain replacement blades, organizational
level. Based on mean MMH per replacement blade.

C req ■ < R mil> * < M req> * B req

(69)

4.10 Cost of blade installation, organizational level. Based on mean
MMH per blade installation.

^nst " < R mil> * ( M inst> * B ri

(70)

4.11 Cost to dispose of scrap, organizational level. Based on mean
MMH per blade scrappage.

CS 0 = < R mil> * < MS o> * BS o

(71)

4.12 Cost of shipping preparation, organizational level. Based on mean
MMH per shipped blade.

CP Q = (CSHP 0 ) * BO dep (72)

4. 13 Blade contribution to maintenance cost, organizational level.

CM 0 = Q on + CR 0 n + C rem + Q 0 ff

+ CR off + c req + Qnst + ^o + CP o

(73)

4.14 Cost of blade inspection, direct support level. Based on mean
MMH per blade received,

^ds ~ ( R mil) * (M^ds) * B ds

(74)

4.15 Cost of blade repairs, direct support level. Based on a s ecified
mean MMH and material cost per blade repair.

CR ds =

(R mil ) * (MR ds ) + (CMR ds )j * BR ds < 75 )

4.16 Cost to dispose of scrap, direct support level. Based on mean
MMH per blade scrappage.

CSds = < R mil> * < MS ds> * BS ds

(76)

4.17 Cost of shipping preparation, direct support level. Based on mean
MMH per shipped blade.

CP ds = <CSHP ds ) * BDS dep

(77)

227

4.18 Blade contribution to maintenance cost, direct support level.

^ds = a ds + CR ds + ^ds + CP ds (78)

4.19 Cost of shipping blades to depot. Based on cost per shipped blade.

C^dep " (CSH US ) * B^gp (79)

4. 20 Cost of blade receiving and inspection, depot level. Based on
MMH per blade received.

^dep = ( R civ) * (^dep) * p dep (80)

4. 21 Cost of blade overhauls, depot level. Based on a specified mean
cost per blade overhaul.

C R dep (CQdep) * PR dep

(81)

4. 22 Cost to dispose of scrap, depot level. Based on mean MMH per
blade scrappage.

C^dep - ( R civ) * ( M ^dep) * 8S dep (82)

4. 23 Cost of shipping preparation, depot level. Based on mean
preparation cost per shipped blade.

C p dep = (CSHP dep ) * BR dep (83)

4. 24 Cost of shipping overhauled blades to field from depot. Based on
mean cost per shipped blade.

CSHF dep = (CSH fld ) * BRdep

4. 25 Blade contribution to maintenance cost, depot level.

^^dep ” C^dep + ^dep + ^ R dep

+ ^dep + ^ P dep + ^^ p dep

4. 26 Total blade contribution to maintenance cost, all levels.

CM = CM^ + CM dg + CM r

(84)

(85)

‘o

dep

( 86 )

4. 27 Replenishment GSE cost,organizational level. Based on specified
mean GSE cost per repair and mean GSE support cost per aircraft,

CG 0 = (CGR on )^*^^ )r ^+ (CGR 0 ff) * BR 0 ff (87)

228

MW0FV 'tV*vt*awtrwJ**tv

4.28 Replenishment GSE cost, direct support level. Based on specified
mean GSE cost per repair and mean GSE support cost per aircraft

CG d = (CGRjg) * BR ds + (CGS ds )

( 88 )

4.29 Replenishment GSE cost, depot level. Based on specified mean
GSE cost per repair and mean GSE support cost per aircraft.

CG dep = (CGR dep ) * BR dep + (CGS dep ) (g9)

4. 30 Total blade contribution to replenishment GSE cost, all levels.

CG = CG q + GG^g + CG^gp (90)

4. 31 Blade life-cycle cost. Blade contribution to aircraft life-cycle cost

BLCC = C fly + ^gp + C rsp + CM + CG ( 91 )

229

5. 0 AIRCRAFT COST EFFECTIVENESS

5.1 Aircraft fuel and oil cost. Based on average mission fuel flow and
cost per pound of fuel.

CPOL = (Quel) * (FF) * (L a ) (92)

5.2 UH-1 nonvariable life-cycle cost. UH-1 life-cycle cost less UH-1
POL and blade life-cycle costs.

LCCn V = (LCC uh ) “ (CPOL uh ) - (BLCC UH ) (93)

5.3 Aircraft life-cycle cost. Based on UH-1 nonvariable life-cycle
cost plus the candidate system POL and blade life-cycle costs.

LCC = LCC nv + CPOL + BLCC (94)

5.4 Aircraft cost effectiveness - ton-knots/$ . The ratio of mission
effectiveness to life-cycle cost.

E ce= fee < 95)

5.5 Baseline UH-1 fleet effectiveness. The mission effectiveness of a
fleet of 1000 UH-1 aircraft.

fe uh = 1000 * (E mUH ) (96)

5. 6 Fleet effective cost. The life-cycle cost of a fleet of candidate
aircraft where fleet size is adjusted to maintain the baseline fleet
effectiveness of 1000 UH-1 aircraft.

FEC =

(97)

230

6.0 MISSION ANALYSIS

Improvements in blade producibility and repairability to reduce life —
cycle cost may penalize weight or aerodynamic efficiency. These pen¬
alties are acceptable if overall cost effectiveness improves.

The impact of a blade design change on mission effectiveness depends on
the requirements of the particular mission. For example, additional
blade weight is a significant penalty only when gross weight is a mission
constraint. Similarly, a penalty in rotor figure of merit will not be
serious for a mission which is never power limited.

The utility role of the UH-1 demands that it operate throughout a wide
range of conditions. This usage cannot be accurately represented by a
single arbitrary design mission. To handle this situation, a Sikorsky
simulation program was used to operate the UH-1 in a probabilistic
mission environment to establish average overall productivity. One
thousand individual mission sorties were simulated for each candidate
blade configuration.

The mission environment used in the simulation was defined by proba¬
bility distributions of the following parameters:

1. Takeoff pressure altitude

2. Ambient sea level air temperature (standard altitude lapse

rate was assumed)

3. Required payload

4. Sortie radius

5. Percent of outbound payload carried inbound

6 . Cover time per sortie

7. Takeoff hover power margin (fraction of HOGE power actually

required)

8 . Cruise elevation above takeoff

Probability distributions are shown in Figure 85. Altitude and tempera¬
ture variations are taken from Reference 15. Sortie radius is distributed
about a mean of 25 nautical miles. Takeoff power margin is based on
zero-wind vertical takeoff being required 20% of the time. At the other
extreme, favorable terrain or wind conditions are assumed to allow
operation at higher weights than provided by adherence to the UH-1
cockpit placard criterion. This criterion, discussed in Reference 15,
calls for a 3% Ni margin at 2-foot skid height, and corresponds to about
90% of HOGE power. Required payload is a demand function indepen¬
dent of capability. It averages 2150 pounds and exceeds the internal
loading limit of 2420 pounds for 10% of the time. Cruise elevation
averages 1500 feet above takeoff as defined in Reference 16.

231

Other inputs to the mission analysis include UH-1H rotor parameters,
engine performance, parasite drag, basic operating weight, and con¬
straints imposed by drive system rating, structural design gross weight,
fuel capacity, and component life allowable speed envelope. Based on
Reference 17, UH-1H parameters are:

Rotor Diameter: 48 ft
Total Blade Area: 84 ft 2
Basic Operating Weight: 5387 lb

Engine: T53-L-13 rated at 1400 Military hp at SLS, with altitude/
temperature and SFC performance per Lycoming Spec.

104. 33

Drive System: 1100 hp flat rated
Red Line Speed: 130 kt indicated
Maximum Gross Weight: 9500 lb

Component Life Allowable Speed Envelope: References 17 and 18

Simulation of the UH-1H with existing blades in the established mission
environment yielded the following results:

Average Takeoff Gross Weight: 7986 lb
Average Outbound Payload: 2051 lb
Average Cruise Speed: 120 knots true
Average Fuel Flow: 533 lb/flight/hr
Average Sortie Flight Time: 0.52 hr

Average Productivity: 50. 3 ton-knots (outbound payload times
sortie radius over sortie time)

Percent of time takeoff limited by gross weight capability: 3.5%
Percent of time takeoff limited by available power: 1. 6%

Percent of time cruise limited by available power: 0.1%

Percent of time cruise limited by component life speed
envelope: 99. 9 %

232

Probability

Takeoff Altitude

Sea Level Temperature

Required Payload ~ 1000 LI) Sortie Radius N.Mi.

.8] Percent of Outboun d
Payload

— ! Carried Inbound

.4-1 :-

.4-

Hover Time
Per Sortie

°0 40 80 120 “0 .2 .4

Percent of Outbound Payload Hover Time

Fraction of Hover OGE Power A Elevation 1000 Ft
Figure 85. UH-1H Mission Environment.

233

APPENDIX IV
UH-1H BLADE DATA

TABLE XXXXIV. UH-1 ROTOR BLADE DESIGN COST COMPARISONS

The following cost model values were supplied by the Government to
standardize the various rotor blade comparisons. The current UH-1
rotor blade values are listed, together with values of the candidate blade
that were considered relatively insensitive to variations in design. Where
values of the candidate blade were not supplied, they were developed by
the Contractor for use after approval by the Government Contracting
Officer.

Current UH-1

Candidate

Blade Life Hours

2500

Aircraft Life Hours

5000

Same

Aircraft Fleet Size

500-1000-2000

Same

Aircraft Attrition

Zero

Same

Blade Set Attrition

.0003/Flight Hr

Same

Time of Blade Initiation

Original Production

Same

Cost of One Blade

$3000

Experience Curve Position

10,000 Blades

Same

;Blade Spares Inventory

30% of Installed

j% Inherent Damage

29.2%

“

|% External Damage

70.8%

Blade Time Between Inherent Damage

547 Hours

! Blade Time Between External Damage

400 Hours

Same

Repair Performance Degradation

Zero

Same

|Cost Field, Org Mil Labor /Hr

4.00

Same

% Military Labor, Field

100 %

Same

Field Overhead and Support Cost

Zero

Same

MMH Each Blade Removal

3.75

Same

MMH Disposition, Inspect

1.5

Same

MMH Repair, Field

“

Parts Material Cost/Repair (Fid)

$5.00

GSE, Tooling Cost/Repair (Fid)

Zero

MMH Obtain Replacement Blade

3.0

Same

MMH Ops, Inventory, Requisition

3.0

Same

MMH Blade Installation

3.75

Same

j% Field Repairs Require Removal

100 %

% Removed Blades Scrapped, Org

30%

—

% Removed Blades Repaired, Org

12 %

!% Removed Blades to Depot Repair

58%

% Depot Received Blades Scrapped

88 %

% Depot Received Blades Overhauled

32%

Shipping, 8000 Mi, Surface, Blade

$90

Same

234

TABLE XXXXIV.

Continued

Current UH-1

Candidate

Shipping, 8000 Mi, Surface, M-T

545

Same

Container

Rotor Blade Container, Reuseable

$200

Same

Preparation for Shipping, Field

$70

Same

% Surface Shipping to CONUS

100 %

Same

% Mil Air Shipping from CONUS

100 %

Same

8000 Mi Mil Air Shipping

$130

Same

% Civilian Labor, Depot

100 %

Same

Composite Civilian Labor Cost, Hr

$12

Same

Blade Overhaul Cost, Depot

$925

Depot Overhaul and Support Cost

Zero

Same

MMH Receive, Unpack Depot

1.0

Same

MMH Inspect (100% of Rec'd), Depot

1.5

Same

MMH to Dispose of Scrap, E)epot

Same

Preparation for Shipping, Depot

$70

Same

Shipping Containers Required

30% of Installed

NOTES:

(a) Develop R&D, prototype and production candidate blade costs,
determine learning curve equation, assume previous production
of 10,000 units and establish cost at 10,000 unit for use in cost
model and comparison with current iJH-1 blade.

(b) Conduct three separate cost runs for each fleet size, 500-1000 -

2000.

(d) Zero aircraft attrition permits the fleet size to remain constant
throughout the analyses; replacing the blade sets at a rate of
.0003/flight hour accounts for the new set of blades required as
a result of attrition.

(e) External damage is further characterized by the following rates:

Battle Damage 16.0%

Dent 25.4%

Foreign Object Damage 16.0%

Puncture 18.8%

Tear 8.0%

Overstress 15.8%

235

APPENDIX V

COST EFFECTIVE COMPARISON USING MTBR OF 1063 HOURS
SUMMARY

Appendix V has been included to provide the R/M tabulations and the cost
effective studies for Configuration V utilizing 1063 hours MTBR instead of
914 hours MTBR for the UH-i. With the higher MTBR, Configuration V
saves $24. 90 x 10 in fleet effective cost compared to the 1980 baseline
blade. This represents only 5% less than the $26. 22 x 10 6 saved with the
original MTBR (page 249). This appendix also includes the rationale for
use of ton-knots instead of ton-miles

EQUIVALENCE OF MISSION TON-KNOTS AND LIFETIME TON-MILES

Mission ton-knots and lifetime ton-miles are equivalent measures of
effectiveness. We have used mission ton-knots because its smaller
magnitude is more convenient.

UH-1 effectiveness can be measured by total work performed. This work
is expressed in ton-miles. However, work per mission is not an accu¬
rate measure since the number of lifetime missions varies with average
mission time. The faster the average mission, the more lifetime work
is delivered in a given useful life. For this reason, ton-miles per
mission cannot be used.

Although mission ton-miles is not a valid measure of effectiveness,
mission ton-miles per hour, or ton-knots, is, since it is equivalent to
lifetime ton-miles. This equivalence can be illustrated with a simple
example. Consider two helicopters, each capable of carrying 1000 pounds
of payload for 20 miles under average mission conditions. One cruises
at 100 knots, the other at 150 knots. Both deliver 10 ton-miles of work
per mission, and on this basis have equal mission effectiveness. The
faster helicopter, however, can fly more missions in a given 5000-hour
service life. Ignoring mission turnaround time, the 100-knot helicopter
delivers a lifetime work of 250,000 ton-miles (1/2 ton x 20 miles x 5000
hr life/. 20 hr mission time). The 150-knot helicopter delivers 50% more
work, or 375, 000 ton-miles. This same 50% superiority for the faster
helicopter is identified by comparing relative mission ton-knots, 75
versus 50. The smaller magnitude of the ton-knot values makes them a
more convenient way to express overall effectiveness.

236

RELIABILITY ANALYSIS

Reliability analysis was performed to compare the baseline UH-1 Blade
with Configuration V (timeframe 1980) using 1063 hours instead of 914
hours MTBR.

BASELINE MATH MODEL INPUT VARIABLES

Math model input variables for the baseline UH-1 blade were held con¬
stant with the exception of the MTBR used in the basic study. The MTBR
value was changed from 914 hours to 1063 hours. The new MTBR of 1063
hours was apportioned to the external and inherent failure rates as
follows:

Inherent . 000230 = 4, 347 hours
External . 000710 = 1,408 hours
Total . 000940 = 1, 06? hours

These new values were run in the math model to establish a new UH-1
baseline cost effectiveness value.

CONFIGURATION V MATH MODEL INPUT VARIABLES

New math model input variables were developed for Configuration V can¬
didate blade using the reliability apportionment of Table XXXXV, the
reliability analysis of Table XXXVI and the repairability analysis of
Table XXXXVII, based upon the new 1063 hr MTBR for the baseline blade.
The new math model R/M inputs are tabulated in Table XXXXVIII. These
values were then run through the math model to establish the comparative
cost effectivess of Configuration V with the UH-1.

237

TABLE XXXXV. RELIABILITY APPORTIONMENT-BASELINE UH-1

WITH 1063 HOUR MTBR

I. Inherent Damage

Frequency of
Occurrence per

Blade Component

Failure Mode

10& Blade Hours

1. Spar

A. Bonding separates

from core

3.0

B. Elongation of

bushing holes

9.0

C. Cracks

D. Abrasion strip

11.0

separation

16.0

E. Corrosion

F. Pitted, abraded or

10.0

eroded abrasion
strip

10.0

59.0

2. Core (Aluminum)

A. Bonding voids

18.0

B. Water contamination

9.0

27.0

3. Skin (Aluminum)

A. Unbonded at leading

or trailing edge

8.0

B. Corrosion

2.0

C. Cracks

27.0

37.0

4. Retention Bushings

A. Cracks

9.0

B. Wear

8.0

C. Corrosion

2.0

19.0

5. Doublers (includes

A. Bonding separation

4.0

grip and drag plates)

B. Corrosion

2.0

C. Cracks

8.0

14.0

238

TABLE XXXXV. (Continued)

I. Inherent Damage - (Continued)

Blade Component

Failure Mode

Frequency of
Occurrence per
IQ 6 Blade Hours

6. Trailing Edge Strip

Bonding Separation 2. 0

(Aluminum)

Cracks

7.0

9.0

7. Trim Tab

Loose Rivets

1.0

Unbonded

3.0

4.0

8. Counterweights

Loose

1.0

Corroded

2.0

9. General

59.0

Total Inherent Damage

230.0

II. Total External Damage

710.0

III. Total Blade Damage

940.0

239

TABLE XXXXVI. RELIABILITY ANALYSIS - CONFIGURATION V COMPARED TO BASELINE

UH-1 BLADE WITH 1063 HOUR MTBR

o o o

o o

o o o o o

o o o

-4 cn co

as co

io N" -4 co

• • •
O 00 CO

CO CO sO

—l CO CO CN t}*

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o .5 W) CO W)

« £ -j- <u .5

2-2 S S-g

5 JK ft a o
< Q U cq cq

CN CO N 1 lO

TABLE XXXXVI. (Continued)
Externally Caused

O O O O iO o o oooooooooo

o o o o

r-- cn co m o co

CVJ -'f <N CO

'OH^OOv

^ CD T3
CD a CD

B.S-S

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bb TO 3
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h o to ■i 1 m vO n

-ICNCO'tinvONoOOO M pi CO ’t

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H 0 .s

wi {3

CO S

r* U Q-

.5 co w

a.2

a «-i

73 «

73 73 O

CO*

242

243

TABLE XXXXVII. REPAIR ABILITY ANALYSIS - CONFIGURATION V COMPARED TO BASELINE

UH-1 BLADE WITH 1063 HOUR MTBR

>ublers 1. Bonding separation

2. Cracks

3. Corrosion

TABLE XXXXVII. (Continued)

co 0) 'rt
J « 8“

o J cs

(D O CC

T3 9* U
CO iU O

^Qw

o o o o

o o

o o o o

o o

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*-H

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CO *-H

• • • •
cs co Tt<

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cd -h
CO

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Tj" <—• co

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• • • •
CM ()\ O h

—< CM CM

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oS «* i

o o o m

• • • •

<N CO —( O

c) a (u ^ d) u

fc a z dh

w <D T3

a> a. a)

563

bO « X3 5
•s? k V u>
QJ O ^ o

£ P O C

O U 3
p_ o Z CL,

-a

- TD cn
c 0) Od
0) r^H

4_, M M -3
c o u' t;
(U 1-1 3 r Q ,
cq cq ca O (

•—i CN CO ^ 1/3 \C I s - -H CM CO •'T lO 0 -HMOO’tlfi'Oh.oO

bO , 0)

C £ bO

a .8-3

rt 1/2 CD

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L . i-H

1 0*0 a

c Mrtt/l

L a>-a a
.5 bo rt 00
M T3 <D 4^

W W J o

246

Dented

Foreign object damage

TABLE XXXXVIII. MATH MODEL R/M INPUT VARIABLES -

CONFIGURATION V COMPARED TO BASELINE
UH-1 BLADE WITH 1063 HOUR MTBR

Variable Value

1. Aircraft Down Hours

2. Aircraft Aborting Failure Rate

j 3. Blade Mean Time Between In¬
herent Damage
| 4. Blade Retirement Life
1 5. % Damage Repaired On Air¬

craft, ORG Level

6. % Damage Repaired Off

I Aircraft, ORG Level

J 7. % Removed Blades Scrapped

1 at ORG Level

j 8. % Removed Blades Sent to

Direct Support

9. % Received Blades Repaired
at Direct Support

10. % Received Blades Scrapped
at Direct Support

11. % Received Blades Repaired
at Depot

12. Maintenance Man-Hours to
Inspect On Aircraft (ORG)

13. Maintenance Man-Hours to
Repair On Aircraft (ORG)

14. Maintenance Man-Hours to
Repair Off Aircraft (ORG)

15. Maintenance Man-Hours per
Blade Repair (Direct Support)

16. GSE Cost Per Repair (Direct
Support)

17. GSE Cost Per Aircraft (Direct
Support)

18. Parts/Material Cost (Direct
Support)

19. Blade Overhaul Cost (Depot)

4. 3839 Down hours per flight
hour

. 015 Aborting failures per flight
hour

1,397 Blade hours
5,000 Blade hours

1. 0 Percent

0 Percent

12. 0 Percent

88. 0 Percent

87. 0 Percent

10. 0 Percent

55.0 Percent

. 25 Maintenance man-hours
1.0 Maintenance man-hours
0.0

7.5 Maintenance man-hours
$84. 10 per repair
$637. 66 per aircraft
$ 66.00

$822. 00 per blade

248

COST-EFFECTIVENESS EVALUATION

The revised MTBR criterion - from 914 to 1063 hours for the baseline
UH-1 blade - does not significantly alter the s r udy conclusions. The
increase in MTBR reduces the number of damaged blades by about 14%,
so the benefits of blade repairability are reduces slightly. True blade
expendability, where the cost of repairing damaged blades at depot is
greater than the cost of replacing them with new blades, is still not
achieved.

With the higher MTBR the Configuration V blade design saves $24. 90
million in fleet effective cost compared to the 1980 baseline blade. This
is only 5% less than the $26.22 million saved with the original MTBR.

The cost of replacing a damaged blade with a new one is still higher than
the cost of repairing it at depot (these costs are unaffected by MTBR), so
true expendability is still not achieved. The Configuration V blade can be
considered more expendable than before, however, since the fewer num¬
ber of damaged blades means that elimination of all depot repair incurs a
net fleet cost penalty of only $33,000 compared to $66,000 with the origin¬
al MTBR.

Figure 86 compares the sensitivity of aircraft cost effectiveness to blade
acquisition cost for the two MTBR criteria. The impact on both the 1980
baseline blade and the Configuration V blade is shown. Overall cost
effectiveness is slightly improved by the increased MTBR, but the rela¬
tive position of the two blade configurations remains about the same.
Figure 87 shows the cost effectiveness blade acquisition cost trends for
the new MTBR criterion with and without depot repair. These trends can
be compared to those in Figure 53 for the original MTBR criterion.

The impact of the increase in MTBR on the benefits provided by the Con¬
figuration V blade design is summarized as follows:

Improvement relative to 1980 baseline blade

MTBR = 914

MTBR = 1063

Blade life-cycle cost
Aircraft life-cycle cost
(including fuel)

Fleet effective cost

- $27,669

- $27,680

- $26,220,000

- $26,342

- $26,352

- $24,900,000

Tables XLIX and L present the detailed cost effectiveness information
for the 1980 baseline and Configuration V blades, respectively, under the
new MTBR criterion. These compare to Tables XXI and XXIX for the
same blades under the original MTBR criterion.

249

Blade Acquisition Cost - $1000

Figure 86. Impact of MTBR Criteria on Cost Effectiveness

Acquisition Cost Sensitivity.

Blade Acquisition Cost- $1000

Figure 87. Impact of Blade Acquisition Cost - Baseline - 1063 MTBR.

TABLE XLIX. COST EFFECTIVE SUMMARY - 1063 MTBR
BASELINE CONFIGURATION - 1980

Aircraft Mission Effectiveness

Aircraft Life-Cycle Cost

Aircraft Cost Effectiveness

37.466 ton-knots
$1,601,476

23. 395 ton-knots/mega $

Fleet Effective Cost

Fleet size adjusted to maintain
fleet effectiveness of 1000 base¬
line UH-1 aircraft

1,601.48 mega $

Life-Cycle Fuel and Oil Cost

$53,344

Blade Contribution To:

Flyaway cost

Initial spares cost

Replenishment spares cost
Organizational level maintenance cost
Direct support level maintenance cost .
Depot level maintenance cost
Replenishment GSE cost

$ 8,417

$ 2,771

$ 46,023

$ 840

$ 523

$ 3,379

$ 0

Blade Life-Cycle Cost

$ 61,953

Life-Cycle Blades:

Damaged

Repaired at the organizational level
Repaired at the direct support level
Repaired at the depot level

Retired on schedule

Replenished by new spares

9. 40 Blades

0 "

1.13 "

1.75 "

0.97 ”

10.50 "

Expendability

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23.349 ton-knots/megadollars

Fleet effective cost 1604.64 megadollars

252

MV'tujiMi mi. «>u.i

TABLE L . GOST EFFECTIVF SUMMARY

- 1063 MTBR

CONFIGURATION V - 1980

Aircraft Mission Effectiveness

37.431 ton-knots

Aircraft Life-Cycle Cost

$1,575,124

Aircraft Cost Effectiveness

23.764 ton-knots/mega $

Fleet Effective Cost

1,576.58

mega $

Fleet size adjusted to maintain
fleet effectiveness of 1,000 base¬
line UH-1 aircraft

Life-Cycle Fuel and Oil Cost

$53,336

Blade Contribution To:

Flyaway cost

6,668

Initial spares cost

1,371

Replenishment spares cost

22,482

Organizational level maintenance cost

1,195

Direct support level maintenance cost

1,555

Depot level maintenance cost

329

Replenishment GSE cost

2,011

Blade Life-Cycle Cost

35,611

Life-Cycle Blades:

Damaged

14. 26 Blades

Repaired at the organizational level

0.14 "

Repaired at the direct support level

10.81 "

Repaired at the depot level

0.20 "

Retired on schedule

0.30 "

Replenished by new spares

6.41 "

Expendability

Scrapping blades normally sent to depot
for overhaul yields the following:

Aircraft cost effectiveness 23. 759 ton-knots/megadollars

Fleet effective cost 1576. 91 megadollars

{

253

APPENDIX VI

PLAN FOR FUTURE HARDWARE EVALUATION

INTRODUCTION

In compliance with the requirements of Contract DAAJ02-71-00046, this
appendix presents the plan for a hardware evaluation of an expendable
UH-1H blade. In view of Sikorsky Aircraft's experience and test facili¬
ties as related to articulated rotors, it is more practical to evaluate
expendable blade concepts on the Sikorsky S-61 helicopter. This is con¬
sidered reasonable because the expendable concept should be applied to
any helicopter application. A proposal to evaluate an expendable S-61
main rotor blade can be submitted upon request.

The selected design was Configuration V, which was found to be the most
cost-effective blade for 1980. This design consists of an all composite
twin-beam structure fabricated in two half sections. A development pro¬
gram, prior to the Plan for Future Hardware, would be required to
develop the advanced process for fabrication.

This appendix includes the development program for the pultrusion pro¬
cess, blade design and fabrication, ground tests, whirl tests, flight tests
and operational suitability for field service of Configu-ruion y. Costs
and schedule are also included for the Plan for Future hardware.

DEVELOPMENT PROGRAM

Before the Plan for Future Hardware Evabution is started, a develop¬
ment period is required for the pultrusion process.

The first phese would be to fabricate simple flat sheets to determine if
the . 020 inch to . 030 inch bias cloth on the trailing edge truss of Con¬
figuration V can be processed successfully with present dies. Problems
may arise with die fouling or freezing which may require die modifi¬
cations or resin system changes. Two investigations will be made of
fabrication, one for material @ ±45° orientation and another with conven¬
tional 90° cross weave. Each type would be fabricated for test evaluation.
Sufficient material would be run through the experimental die to give an
indication of possible future production problems, such as progressive
die fouling, unpredictable jamming or other complications, so that these
can be properly taken into account before proceeding.

The second phase would be to proceed with tooling for a partial section
of the truss fairing. This step would be used to check out the ability to
flow the skins together properly in the final webbed shape selected, to

254

determine what warp or distortion problems occur, and to provide sample
sections for test evaluation.

The third phase would be to produce the final production fairing section.
This would involve the design and procurement of a full-scale die con¬
figuration suitable for production. A number of trailing edge fairings
would be produced for test evaluation using this die. There would be
some flexibility in the tooling to allow for changes in wall thicknesses,
material composition, and distribution of webs.

The fourth phase would be an expansion of Phases I through III, i. e., the
fabrication of the twin-beam spar. The spar beam is a solid section and
will not be as complex as the web section. However, combining this
component with the truss will require additional dies and mandrels and
some development. Several sections would be fabricated and then sub¬
jected to quality control for inspection of dimensional tolerances,
straightness, bowing, contouring, etc. Several specimens would also be
evaluated by structural testing. Work would continue during this phase
until the parts can meet the minimum specifications of strength and
dimensional requirements. This development time for fabrication of
sections and subsequent structural test evaluation should be approxi¬
mately 1-1/2 years.

PLAN FOR FUTURE HARDWARE

The plan is for fabricating a total of 8 full blade assemblies oi Configu¬
ration V. Four of these assemblies will be used for fatigue structural
tests. Each one of the four will serve as two test specimens, one in¬
board and one outbot rd, resulting in a total of four inboard and four out¬
board test specimens. A total of four full blade assemblies will be
utilized for whirl and flight tests.

In addition, an equal number of UH-1 blades will be subjected to the same
structural, whirl and flight tests to obtain a valid comparison with the
new blade.

Figure 88 shows the phases of the program which would cover a period
of two years. The cost breakdown through flight test is shown below.

Cost Breakdown

Engineering - - - - $800,000

Manufacturing-$884,000

Materials and -$116,000

Direct Cost _ _

$1,800,000

255

Figure 88. Plan for Future Hardware Evaluation.

DESIGN AND FABRICATIO N

Detail and assembly drawings will be prepared and released to
Purchasing and Manufacturing for procurement of materials and fabri¬
cation of in-house portions of the blade. In addition, manufacturing
and design engineers will support the subcontractor responsible for the
fabrication of the blade half section by the pultrusion process. Design
and material specifications will be developed to provide high quality
components.

The upper and lower tool molds will be machined from aluminum cast¬
ings. Machining will be accomplished on a tape controlled milling
machine and the finished tool surface will be machined to within . 005
inch of the required aerodynamic contour. The root end of the mold
would be contoured to accommodate the enlarged attachment area of the
blade. Dies for the root doublers and forging drag plates would be
procured for fabrication and for later assembly onto the blade.

The half blade pultrusion assembly (which includes the fiberglass spar,
trailing edge truss with combinations of carbon and fiberglass, the lead¬
ing and trailing edge carbon doubler and spline) would be inspected by
quality control upon completion of each part. A short section would be
cut from each end for examination by Materials and Processing personnel
and for additional test evaluation. The mass distribution of each half
would also be prechecked prior to applying counterweights to minimize
any balancing problems which may occur after assembly.

The finished root end doublers and drag plate would be placed in their
respective mold,and each half blade section would be placed in position
over the doublers. The leading edge counterweight consisting of
elastomer containing lead shot would be cast in place using a retaining
tool. The honeycomb core and the foam-in-place would next be assem¬
bled in the blade. A routing tool which fits on the mold will then be used
to cut the assembly to the chordline. The mass distribution of the two
machined blade halves will then be determined and corrections made to
the leading edge counterweights.

The two blade halves will then be joined by structural adhesive. A
polyurethane erosion coating will then be applied to the leading edge,
and tip weights will then be added to adjust the blade spanwise balance.
The final blade assembly will be inspected using ultrasonic
coin tap, visual, and dynamic balance techniques.

Throughout the blade fabrication process, Engineering and Manufactur¬
ing Engineering will revise and update the manufacturing operations
plan. At the completion of the fabrication effort, a complete set of

257

operation sheets, with all modifications, will be compiled and evaluated
to establish a firm manufacturing base.

STRUCTURAL TEST PLAN

Full-scale blade specimens of Configuration V and specimens of the
present UH-1 baseline production blade will be laboratory tested to pro¬
vide comparison of the expendable and production blades and to verify
structural integrity for flight testing of the expendable blade. This is
accomplished by over stress fatigue testing of these full-scale blade
specimens to

1. Establish mean strength

2. Verify rotor blade design analysis

3. Determine rotor blade failure modes

4. Determine rotor blade fail-safe characteristics

Overstress fatigue testing is conducted on eight representative speci¬
mens of both the Sikorsky Aircraft designed expendable main rotor blade
and the production UH-1 main rotor blade ; four outboard specimens of
both configurations, representing the most highly stressed outboard
blade section, and four inboard specimens of both configurations, rep¬
resenting the hub-to-blade attachment area or root end.

The specimens are fabricated from four full-scale blades of both con¬
figurations by separating the test portions from the full-scale blades and
modifying the specimen ends to accept load fittings compatible with the
fatigue testing equipment. See Figures 89 and 90.

Both configurations are tested to avoid interpretation of any differences
in mean strength which may occur, as a result of differences in test
techniques, differences in handling of mean strength data, of systematic
error, i. e., a difference in the mean strength between blades tested
now and blades tested in the original substantiation several years ago.

Overstress relates the blade specimen test stress level to the aircraft
blade operating stress level. The blade specimen test level is higher
than the analytically predicted blade mean strength stress level and is
much higher than the blade operating stress level. Testing at an over¬
stress level determines the mean strength of the blade in a reasonable
time frame by initiating a crack or accumulating the number of cycles

258

1 Ft (typ)

259

Figure 89. Outboard Specimen.

chat is predetermined to be a runout. If a runout occurs, subsequent
test specimens will be tested at a higher stress level tc initiate a crack.
If the specimen does not separate, crack propagation testing is conducted
by applying stresses representative of normal aircraft blade operating
conditions and measuring crack length. Crack propagation testing of
the specimen shall be discontinued after 25 hours.

The outboard specimen is tested as a resonant pin-pin beam. An axial
load is mechanically applied to simulate centrifugal force. By orient¬
ing the specimen at a particular angle with relation to the plane of the
pins that support the specimen, blade edgewise and flatwise stresses
are obtained. The specimen is mechanically forced to vibrate at a
frequency close to its natural frequency, resulting in vibratory stress
caused by deflection of the specimen. The closer the forcing frequency
is to the natural frequency of the specimen, the greater the beam de¬
flection and resulting stress. Required stress levels are obtained by
increasing the forcing frequency.

The root end specimen is tested as a cantilevered beam. An axial load,
mechanically applied, simulates centrifugal force. By orienting the
specimen at a particular angle with relation to the direction of the
applied load, blade edgewise and flatwise stresses are obtained.

Both the outboard and the root end specimens are instrumented in
selected areas to measure strains. Edgewise and flatwise strains will
be measured.

Each specimen is tested at a constant stress level, and strength data is
presented in the form of a stress-cycle (S-N) plot. See Figure 91.

An S-N curve shape for the applicable material is drawn through the
test data points. Aircraft blade operating stresses are related to this
curve to determine the structural reliability of the blade. See Figure 92.

Crack propagation data will be presented in the form of a real time plot
of stress vs. aircraft flight time. See Figure 93.

A static rap test establishes the edgewise and flatwise natural frequency
of the blade. A strain-gaged blade is hung vert ically and struck with a
mallet in each of the three (edgewise, flatwise, and torsional) directions.
The output of the respective strain gages determines the three respective
natural frequencies. See Figure 94.

ROTOR SYSTEM WHIRL TESTS

Comparison rotor whirl tests of the expendable and standard UH-1 blades

261

Figure 91. Stress Level of Operation.

Figure 92. S-N Strength Data.

Flight Hours

Figure 93. Crack Propagation.

262

verify equal hover performance, aerodynamic and aeroelastic predic¬
tions of stability, and provide adequate endurance validation for flight
testing. The rotor whirl tests consist of the following:

1. Aerodynamic and dynamic balance adjustments to provide
identical tracking characteristics in flight.

2. Comparative hover performance tests of the expendable blades
and standard UH-1 blades to demonstrate equal performance.

3. Stress and motion surveys to validate design predictions of
stress levels and frequency response.

4 . Thirty hours of endurance at conditions simulating anticipated
flight loads.

5. Dynamic checkout of blade instrumentation prior to flight
testing.

6. A 1-minute rotor overspeed test at 110% of limit power off
rotor speed.

Aerodynamic and dynamic balancing is accomplished on the Sikorsky
2000 HP Main Rotor Test Stand by adjusting the blade pitching moment
and track characteristics alike on the expendable blades. Aerodynamic
balancing consists of adjusting the trailing edge trim tabs to match the
blade pitching moments at low angles of attack. Dynamic balancing
entails matching the blade pitching moments and track at high collective
pitch angles by chordwise adjustments to tip weights. The blade pitch-
moments are obtained by measuring the steady loads in the rotor head
rotating control rods. Track measurements are obtained using a Chicago
Aerial Electronic Blade Tracker.

Comparative hover performance tests on the expendable and standard
UH-1 blades are performed on the 30-foot-high Sikorsky 2,000 HP Main
Rotor Test Stand to demonstrate equal lift capability. Rotor thrust,
power, blade angle and pitching moment data are obtained at tip Mach
numbers corresponding to 90%, 100% and 110% of normal rotor speed for
both blade types from zero thrust to the maximum attainable thrust as
limited by the structural or geometric limits. To keep wind effects to a
minimum, data are obtained when the wind velocity does not exceed 5
knots. Following correction of the data to sea level standard conditions
(59°F, 29.92 inches Hg, and zero wind), the results are presented as
comparison plots for each tip Mach number. Figures 95 and 96 show the
Test Stand and Thrust/Performauce comparisons.

264

I i

j f

• j

CL)

1-4

1-1

qi - 3snjqx

266

Figure 96. Rotor Hover Performance Comparison,
Configuration V vs Standard Blades.

MT.-9

Stress measurements detained throughout the rotor operating range as
functions of rotor speed, rotor thrust,and blade flapping validate design
stress levels and natural frequency response throughout the entire
operating range to verify design criteria. The data are acquired on
magnetic tape to facilitate data reduction and analysis.

Thirty hours of endurance testing at 8000 pounds of thrust and blade
flapping loads simulating anticipated flight loads provides adequate
assurance of structural integrity prior to flight tests.

The final whirl test consists of a 1-minute overspeed run at 110% of
the limit power off to demonstrate safe operation at maximum autorota-
tive rotor speed. Following the whirl tests and prior to flight tests,
the blades are inspected to verify that no defects resulted from the whirl
tests.

267

5-HOUR FLIGHT EVALUATION

A 5-hour flight test will be conducted for hardware evaluation of the
expendable UH-1 main rotor blade. Flight tests of the Sikorsky-designed
expendable main rotor blades will determine the structural airworthiness
of the blades and investigate the effect of the new component on general
aircraft handling and vibration characteristics. Blade stress and motion
data will be obtained throughout the established aircraft operating
envelope with vibration and handling qualities data obtained simultaneous¬
ly with the structural measurements. Principal measurements will also
be obtained on selected main rotor components to provide safety of flight
and to evaluate the influence of the new component on stresses and loads
in other areas of the rotor system. The flight tests will be conducted at
the Sikorsky test facility, Stratford, Connecticut, at altitudes ranging
from ground level to approximately 3,000 feet density altitude. A bailed
UH-1 and pilot will be required for the flight test program.

The test rotor blade will be extensively instrumented using strain gages
(approximately 30) to determine the stress distribution and blade response
characteristics. Chordwise and normal bending stresses will be meas¬
ured at several blade stations including the root end attachment area and
approximately five additional spanwise locations. Stress levels will
also be recorded on the structural trailing edge of the blade and torsional
stresses measured at the 30% and 75% blade radius stations. A typical
Sikorsky main rotor blade strain gaged to recorded edgewise, flatwise
and torsional stresses is illustrated in Figure 97. Additional main
rotor instrumentation will include blade motions, control loads, drag
brace load, shaft bending, and edgewise and flatwise stresses on the
hub/sleeve assembly. Aircraft attitude, control positions, vibration
levels and load factor will also be recorded along with pertinent cock¬
pit data (airspeed, rotor speed, altitude, etc.) which are necessary to
document the flight conditions.

The structural characteristics of the expendable blade will be primarily
evaluated at the maximum aircraft allowable gross weight at both the
forward and aft center-of-gravity extremes. An initial hovering flight
will be conducted at light gross weight followed by subsequent flight
testing at the maximum gross weight condition. The heavy weight test¬
ing will require a gradual buildup in forward speed until maximum for¬
ward speed is achieved. During this phase, selected parameters will be
monitored by telemetry to provide safety of flight. A cursory check of
blade stress will also be conducted at the aircraft basic design gross
weight to verify that the maximum blade stress levels occur at the alter¬
nate gross weight configuration. No extreme altitude or envelope type
tests are planned for this evaluation.

268

Test data will be acquired at maximum gross weight for the follow¬
ing flight conditions:

1. Rotor engagement.

2. Hovering flight including sideward flight, rearward flight,
control reversals, hover turns and rotor speed sweeps.

3. Level flight to V m ax at various rotor speeds.

4. Normal maneuvering and control reversals within the
approved operating flight envelope.

5. Partial power descents at three airspeeds and two rates
of descent.

6. Autorotations at three airspeeds at normal, maximum
and minimum approved power-off rotor speeds.

7. Takeoff, climbs, climbing turns, and final approach and
landing.

The cursory check of blade stresses at the aircraft design gross weight
will be limited to but will not necessarily include all of items 1, 2, 3,
and 7 above.

DESIGN REPORTS

Three technical reports will be prepared for this study: Blade Loads
Report, Stress Report and Test Report. The Blade Loads Report will
summarize all the aerodynamic loads subjected to the blade throughout
the aircraft flight spectrum. The Stress Report will present a detail
stress analysis of all blade components for the most critical aircraft
maneuvers. It will also present a blade life calculation for the flight
spectrum established ir. the Blade Loads Report. The Test Report would
include all the results of structural ground testing and the stability and
handling qualities from flight testing.

270

OPERATIONAL SUITABILITY EVALUATION

A 2-year field service evaluation of the expendable main rotor blade
is required to demonstrate component reliability and maintainability.
Following substantiation of the expendable blade,environmental testing
under actual field service conditions is proposed. Field service
experience will be obtained by installing 6 sets of blades on operational
UH-1H helicopters operating under normal field conditions. Inspection
of the blades would be conducted daily and the findings reported in field
service reports. The reports should specify the operating environment,
effects of erosion by sand, dust, rain, and describe any damage in¬
curred, along with the procedures for repairing the damaged blade. The
damage and repair data should include but not be limited to:

1. Time to damage

2. Type of damage

3. Description of repair procedure

4. Man-hours required for inspection, repair,
and checkout of component

5. Problems encountered with repaired components

The inspection procedure should be continued for 2 years from

time of delivery, with the environment of the blade being varied as much

as possible.

After the first year, an evaluation will be made of the expendable blade
to determine feasibility of incorporation into production. Vigilance
will still be maintained on the blades installed for the 2-year period.

271

CONTINGENCIES

1. A total of eight UH-1H blade assemblies should be supplied for
structural and whirl test evaluation.

2. UH-1H rotor head components and assemolies should be supplied as
required for ground tests.

3. A bailed UH-1 helicopter and qualified test pilot are required for
the program.

4. If a qualified pilot is not provided with the aircraft, Sikorsky test
pilots will be used. However, additional costs will be required to
qualify two Sikorsky pilots at an off-site facility.

5. Sikorsky personnel can support the test aircraft; however,

appropriate manuals and handbooks have to be provided, preferably
before the test program commences.

RECOMMENDATIONS

It is recommended, as a first phase, that a development program for
the pultrusion process of manufacturing a one-piece cover be started
immediately. In addition to supplying the ground work for the 1980 twin
beam concept, the process is also applicable to Configuration VI, the
aluminum spar with the automated cover. It is further recommended
that the expendable blade concept be evaluated on a Sikorsky S-61 helicop¬
ter. This program would evaluate the cost, reliability, repairabilily
and the aeromechanics of the twin-beam expendable blade concept.

272

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Full text of "DTIC AD0762198: Expendable Main Rotor Blade Study"

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